AD-766 307 United Aircraft Corporation · ad usaamrdl technical report 73-11 0vibration effects on...
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o - _ ; - t _ - - " ... . _- L .. . ... _ - AD-766 307 VIBRATION EFFECTS ON HELICOPTER RELI- ABILITY AND MAINTAINABILITY Angelo 2-. Veca United Aircraft Corporation Prepared for: Army Air Mobility Research and Development Laboratory April 1973 DISTRIBUTED BY: WEM National Teconical Information Service U. S. DEPARTMENT OF COMMERCE 5285 Port Royal Road, Springfield Va. 22151 I- _______Jl___IIII L I _
Transcript of AD-766 307 United Aircraft Corporation · ad usaamrdl technical report 73-11 0vibration effects on...
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AD-766 307
VIBRATION EFFECTS ON HELICOPTER RELI-ABILITY AND MAINTAINABILITY
Angelo 2-. Veca
United Aircraft Corporation
Prepared for:
Army Air Mobility Research and DevelopmentLaboratory
April 1973
DISTRIBUTED BY:
WEMNational Teconical Information ServiceU. S. DEPARTMENT OF COMMERCE5285 Port Royal Road, Springfield Va. 22151
I- _______Jl___IIII LI
_
AD
USAAMRDL TECHNICAL REPORT 73-11
0VIBRATION EFFECTS ON HELICOPTERRELIABILITY AND MAINTAINABILITY
By
Angelo C. Veca
April 1973
EUSTIS DIRECTORATEU. S. ARMY AIR MOBILITY RESEARCH AND DEVELOPMENT LABORATORY
FORT EUSTIS, VIRGINIACONTRACT DAAJ02-71-C-0037SIKORSKY AIRCRAFT DIVISION
UNITED AIRCRAFT CORPORATION
STRATFO!?D, CONNECTICUT
Approved for public release;distribution unlimited.
Rop-d.ced by
NATIONAL TECHNICALINFORMATION SERVICE
5 044, VA 221It
DISCLAIMERS
The findings in this report are not to be construed as an official Department of the Armyposition unless so designated by other authorized documents.
When Government drawings, specifications, or other data are used for any purpose otherthan In connection with a definitely related Government procurement operation, the
United States Government thereby incurs no responsibility nor any obligatic. i whatsoever;and the fact that the Government may have formulated, furnished, or in any way suppliedthe said drawings, specifications, or other data is not to be regarded by implication orotherwise as in any manner licensing the holder or any other person or corporation, orconveying any rights or permission, to manufacture, use, or sell any patented inventionthat may in any way be related thereto.
Trade na m es cited in this report do not constitute an official endorsement or approval ofthe use of such commercial hardware or software.
DISPOSITION INSTRUCTIONS
Destroy this report when no longer needed. Do not return it to the originator.
Lp
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UNCLASSIFIEDstculity Classilatim
DOCUMENT CONTROL DATA -R &DIS*e.witv cleam D0hation, of lots, bodt, of .astrot Iend irdo ine arnotatio.. must be entems! whon the ... rall twpoef to r.saviflestI ,ORIGINATING ACTIVITy (COWI~.tO eUtfti)U. REPORT SECURITY CLASSIFICATION
Sikorsky Aircraft Division dnlssfeUnited Aircraft Corporation 21 classifieStratford, Conaecticut
IREPORT TITLE
IVIBPATICN EFFECTS 0:1 HELICOPTERRELIABILITY AND MAINTAINABILITY
4. UESCRIPTIVX NOTES (I'yps of report end inclusive dates)
Final ReportS A4TP4O.S)(Ji8FatM~ middle Inlitial. le*t Rem.)
Angelo C. Veca
6, ALPORT DATE 7a. TOTAL NO. OF PAGES Tb. orO Reps
Apil 1973 /2 . 5e.CON-MACY OR GRANT NO0. 0e. OMIGINA 1 09S McpGR r NUMUIRI)
DMJO02-71--0037 USAAMKDL Technical Report 73-11A6 PROJECT No.
1F32oLDB38 ______________________
C. 9b. DINER RE[PORT NORS) (An1y 61hor numInhors Owl may be eeel&e1O~le report)
IL Sikorsky Aircraft Report SER-61156710. O!STMIIUV*IOH STATCNT'
Approved for public release; distribution unlimited.
11. SUPPLEIIENT1ARY NOTES 12- SPONSORING MILITARY ACTIVITY
Eustis Directorate, U. S. Army AirIMobility Research & DevelopmentLaboratory, Fort Eustis, Virgi*.ia
12. A ""T ACT
In this study, aifferenzea in reliability and maintainability data were examined ontwo groups of USAF 1-3 helicopters with distinctly different vibration characteristics.One H1-3 helicopter group was equipped with the rotor-mounted vibration absorber, adevice which rvduces helicopter vibration induced by the rotor, and a second aircraftgroup did not have the absorber. The aircraft were alike in all other respects.
The analyses performed on these data show a significant reduction in the failure rateand direct mainti-nance for the H-3 helicopters with absorbers and with reduced vibra-tion levels. The overali H-3 helicopter failure rate and corrective maintenance are
freduced by 48% ant' 38.5%, respectively. The average reduction in vibration level was' 5~.%. c ifed-zyz, coats show a siicf'wiuL reldutuiun or approximately
10% for the overal. aircraft. At the subsystem and component levels, the same re duc-tions are shown in almost every case with the exception of certain navigation and
avionics components.
There are at least 2,z military vibration specifications and standards which specifyvibration criteria for design or test of airborne equipment. No obvious conflictswere found in these slecifications, but they are lacking in requirenents which clearlydescribe realistic vib-ation exposure times for the entire helicopter air vehiclesystem and its componerzts.
Few REPLACER! 00 FORM 947S. I JAM 04. WHICH inD D 110111OI1473 ONOIES FO AI UK UNCLASSIFIEDCesfceo
Ieut k-111A&0
UNCLASSIFIEDSecb'ity'Clasmjllcalon
14. LINK A LINK a LINK CKery WORDS
ROL.9 WT ROLIL WT ROLIC WT
Helicopter
VibrationReliabilityMaintainabilityLife-Cycle Cost
UNCLASSIFIED! l4cwlty classincou~n
* I' ' "> I " ~ I n i m " * . , , i r . . . + . I i
UNCIASSIFIED
security Classificatlon14. LINK A LINK 9 LINK C
KEY WORDSROLE WT MOLE1 WT ROLE WT
Helicopter
VibrationReliabilityMaintainabilityLife-Cycle Cost
Secuty Claslsicatuic
Project IF163204DB38
Contract DAAJ02-71-C-0037USAAR1RDL Technical Report 73-11
April 1973
VIBRATION EFFECTS ON HELICOPTER RELIABILITY AND MAINTAINABILITY
I"Final Report
iF
Sikorsky Aircraft Report SER-611567
By
Angelo C. Veca
iPrepared by
Sikorsky Aircraft DivisionUnited Aircraft Corporation
Stratford, Connecticut
for
Eustis DirectorateU. S. ARMY AIR MOBILITY RESEARCH AND DEVELOPMENT LABORATORY
FORT EUSTIS, VIRGINIA
Approved for public relcase;distribution unlimited.
SUMMJARY
This study assesses the effect of helicopter vibration environ.ent onhelicopter subsystem reliability, maintainability, and life-cycle costs,and the adequacy of design and acceptance test specifications applicableto helicopter vibration.
In this study, differences in reliability and maintainab.lity data wereexamined on two groups of USAF H-3 helicopters with distinctly differentvibration characteristics. One H-3 helicopter group was equipped withthe rotor-mounted bifilar vibration absorber, a device which reduceshelicopter vibration induced by the rotor, and P. second aircraft groupdid not have the absorber. The aircraft were alike in all other respects.
The analyses performed on these data show a significant reduction in thefailure rate and direct maintenance fo' the H-3 helicopters with absorbersand with reduced vibration levels. The overall H-3 helicopter failurerate and co,-rctive maintenance are reduced by 48% and 38.5%, respectively.The average reduction in vibration level was 54.3%. Correspondingly,life-cycle costs show a significant reduction of approximately 10% forthe overall aircraft. At the subsystem and component levels, the samereductions are shown in eianost every case with the exception of certainnavigation and avionics components.
There are at least 2:9 military vibration specifications and standardswhich specify vibration criteria for design or test of airborne equipment.No obvious conflicts were found in these specifications, but they arelacking in requirerients which clearly describe realistic vibration exposuretimes for the entire helicopter air vehicle system and its components.
As shown by this study, reduction in vibration levels can significantlyimprove reliability and reduce maintenance and life-cycle costs. Theresults also suggest that the useful life of an aircraft can be extendedbeyond current limits simply by reducing vibration exposure.
iii
FOREWORD
The work for this study was authorized by Contract DAAJ02-71-C-0037, ProjectIIh'6320D.B38, issued by the Eustis Directorate, U. S. Army Air MobilityLResearch ana Development Laboratory, Fort Eustis, Virginia under the tech-nical cognizance of Major A. Gilewicz.
The Sikorsky Aircraft personnel involved in performing or assisting in
this study were:
Dr. David Jenney, Chief of System Engineering Design.
Mr. Miller A. Wachs, Supervisor of Reliability and Maintainability,Project Manager.
Mr. Robert Oaseria, Systems Analyst, Life Cycle Costs.
Mr. James Duh, Loads & Criteria, Vibration Level Mapping andCalculations.
Fr. Michael Starzyk, Standards, Specification Aasessment.
Mr. Spencer Lauer and Mr. Thomas Chernesky, Technical Computing,AFM 66-1 Data Reduction.
pi
I
Preceding page bvii
fTABLE OF CONTENTS
Page
SUMMARY ................................. .1
FOREWORD ........ ...... .............................. v
LIST OF ILLUSTRATIONS ........ ......................... viii
LIST OF TABLES ........ .... ........................... xi
LIST OF SYMBOLS ......... .. ........................... xiii
INTRODUCTION ......... .... ............................ 1
METHOD AND RESULTS ......... ... ........................ 5
Aircraft Population and Data Separation ...... ............. 5Determination of Vibration Magnitudes .... ... .............. 9Reliability and Maintainability Data Source and Analysis ..... 25Individual Aircraft Reliability and MaintainabilityComparison ......... . . . . . .......... ......... 86Life-Cycle Cost Model and Life-Cycle Cost Determination ..... .. 86Vibration Design and Acceptance Test SpecificationAssessment .......... ............................ .. 98
DISCUSSION OF RESULTS ........ ........................... 106
Reliability ......... ... ........................... 106Maintainability .... ..... ......................... ... 208Specification Assessment ..... .. ......................... l1
CONCLUSIONS ..... ...................... ..... ... . 12
RECOMMENDATIONS .......... ... ........................... 113
LITERATURE CITED .... ..... ... .......................... l11
APPENDIXES
I. Climatic Briefs .... ...... .. ............. .... 115II. Excerpts From Assessed Specifications. . ..... ............ 122
DI$TRIBUTION . . . . . . . . . . . . ... . . . . 125
Preceding page blank vii
LIST OF ILLUSTRATIONS
Figure Page
1 Fixed-Wing Electronic Reliability vs Rotary-WingReliability ....... ..... ...................... 2
2 Fixed-Wing Hydraulic Reliability vs Rotary-WingHydraulic Reliability ....... .. .................. 3
3 H-3 Helicopter ....... ... ..................... 6
Bifijar Tuned Vibration Absorber ...... ............ 7
5 Total Vibration Response g Without Absorber ......... ... 13
6 Total Vibration Response g With Absorber ........ 14
7 Extrapolation, Interpolation Vibration g LevelsBL IO(LT) WL 181.5 .............. ................... 16
8 Extrapolation, Interpolation Vibration g LevelsBL !0(RT) WL 181.5 ......... ................... 17
9 Extrapolation, Interpolation Vibration g LevelsSta. 250 and Sta. 300, WL 181.5 ... ............. .... 18
10 Extrapolation, Interpolation Vibration g LevelsBL 0, WL 181.5 ....... ..................... .... 19
11 Linear Extrapolation, Interpolation of Vibrationg Level for Entire CH-3 Aircraft Without VibrationAbsorber ....... ..... ....................... 21
12 Linear Extrapolation, Interpolation of Vibrationg Level for Entire CH-3 Aircraft With VibrationAbsorber ..... .... ........................ 22
13 Vibration Magnitude Profile Without and WithVibration Absorber at Cabin Ceiling Level, WL 181.5 . • 23
14 Vibration Magnitude at Cabin/Cockpit Floor Level,
WL 107....... ................ 24
15 Location Grid for CH-3 Aircraft Components ... ....... 61
16 Comparison of Total Average Failure Rate and MMH/KFHfor Top 13 Aircraft Subsystems ...... ............. 67
17 Comparison of Average Failure Rate and MMH/KFH forSelected Airframe Subsystem Components . ......... ... 68
viii
; "' cl~'
I' [ | " i ' '... . . '... 'll "' i l t# : l 1 v ' , ,. - .. . , - i 2 .... -
Figure Page
18 Comparison of Average Failure Rate and M4H/1KH forSelected Drive Subsystem Components .... ........... ... 69
19 Comparison of Average Failure Rate and M.4H/TH forSelected Utility Subsystem Components ........... 70
20 Comparison of Average Failure Rate and M4H/KFH forSelected Landing Gear Subsystem Components .. ........ .. 71
21 Comparison of Average Failure Rate and I4H/KFH forSelected Fuel Subsystem Components .... ............ ... 72
22 Comparison of Average Failure Rate and 1N4H/'H forSelected Flight Control Subsystem Components ........ ... 73
23 Comparison of Average Failure Rate and N4R/KFH forSelected Cockpit/Fuselage Subsystem Components ... ...... 74
24 Comparison of Average Failure Rate and MMH/KFH forSelected Electrical Subsystem Components .. ......... ... 75
25 Comparison of Average Failurp Rate and MMH/KFH forSelected Hydraulic Power Subsystem Components ......... ... 76
26 Comparison of Average Failure Rate and MMH/KFH forSelected Intercommunications Subsystem Components ....... 77
27 Comparison of Average Failure Rate and MMH/KFH forSelected Radio Navigation Subsystem Components ...... 78
28 Comparison of Average Failure Rate and MOIH/KFH for, Selected Airconditioning/Heating Subsystem Components . . . 79
29 Comparison of Average Failure Rate and W.IH/KFH forall Internal and External Lights ...... ............. 80
30 Comparison of Average Failure Rr.te and MMOH/KFH forall Switches ........... ...................... 81
31 Comparison of Average Failure Rate and MMH/KFH forall Connectors/Plugs/Wiring ..... ................ ... 82
32 Comparison of Average Failure Rate and MMH/KFH forall Hoses and Lines ....... .................... ... 83
33 Comparison of Average Failure Rate and NMH/KFH forall Valves ........... ........................ 84
34 Comparison of Average Failure Rate and MMH/KFH forall Relays .......... ...................... ... 85
ix
Figure Page
35 Distribution of Individual Aircraft Failure Ra-tesComparing Aircraft Groups Without and With Absorber . .. 87
36 Distribution of Individual Aircraft M.i/FH ComparingAircraft Groups Without and With Absorber ... ..... 88
37 Life-Cycle Cost Model ...... ................... ... 89
38 Vibration Specificazion Tree ............... 100
x
LIST OF TABLES
Table Page
I Selected H-3 Aircraft, Locations, and Times .... ........ 8
II H-3 Aircraft Mission Profile ........ .............. 9
III Aircraft Vibration Response ....... ................ 10
IV Total Aircraft System Comparison, Reliability andCorrective Maintenance ...... .................. ... 27
V Component Comparison of Reliability and CorrectiveMaintenance at Subsystem Level ....... ............. 29
VI Compcrnent Comparison of Reliability and CorrectiveMaintenance - Internal and External Lights .... ........ 39
VII Component Comparison of Reliability and CorrectiveMaintenance - Switches ..... .................. .. ..
VIII Component Comparison of Reliability and CorrectiveMaintenance - Connectors/Plugs/Wiring .... ........... 45
IX Component Comparison of Reliability and CorrectiveMaintenance - Hoses/Lines ...... ................. 47
X Component Comparison of Reliability and CorrectiveMaintenance - Valves ...... ................... .... 53
XI Componert Comparison of Reliability and CorrectiveMaintenance - Relays ...... ................... .... 59
XII Individual Aircraft X and MH/KFH Without Absorber ........ 63
XIII Individual Aircraft A and MMH/KFH With Absorber ........ 64
XIV Comparison of Aircraft A and M24H/KFH by PopulationGroup on Aircraft Action Only .... .............. ... 65
XV Comparison of Scheduled Maintenance Actions andMaintenance Man-Hours ....... ................... ... 66
XVI Life-Cycle Cost Model Assumptions .... ............. ... 91
XVII Life-Cycle Savings per Aircraft Resulting FromVibration .Reduction ........ ..................... 94
XVIII Life-Cycle Model Nonbifilar Subsystem Spares andAbort Rates ......... ........................ ... 96
xi
. ,. --,r ,L .. . . : : = " " % " "= " ' ' - ",m" " -#'m = - . . .. . .. . . .. .. .... _ . 4
Table Page
XIX Life-Cycle Savings per Aircraft Resulting FromVibration Reduction ...... .. . ..... . ........... 97
XX Applicable Specifications With Vibration Requirements . 99
XXI Ratio Change in Average Failure Rate and Ratio Change inVibration Level ....... ...................... .. 107
XYII Ratio Change in .2IH/FH and Ratio Change in AverageVibration Level ....... ......................... 110
-XXIII AWS Climatic Brief, Eglin AFB, Fla ................... 115
TXXIV AWS Climatic Brief, Forbes AFB, Kansas ..... .......... 116
XXV AWS Climatic Brief, Parks AFB, P.I ..... ............. 11
XXVI AWS Climatic Brief, Shaw AFB, S. Carolina ..... .. 118
XXVII AWS Climatic Brief, Thule AFB, Greenland .. ......... .. 119
XXVHII AWS Climatic Brief, Wcodbridge Aerodrome, England ....... 120
XXIX Climatological Data for Elmendorf AFB, Alaska .......... 121
xi
xti
-- --
CLIST OF SYMBOLS
A lateral acceleration amplitude
B vertical acceleration amplitude
DH/D down hours per day
F primary rotor excitation frequency, rad per secn
2g gravitational acceleration, 32.2 fT per sec
H vibratory frequency, cycles per secz
LCC life-cycle costs
MI maintenance sensitivity index
P M4H/FH maintenance man-hours per flight hour
MTBF mean time between failures
n number of blades
OD operational day, hr
OPave average operational payload, lb
PL aircraft payload, lb
Rtotal vibration response
R mean square ratio
RI reliability sensitivity index
t mission time, hr
T test statistic
V cruise speed, ktc
Xlateral acceleration
Zvertical acceleration
AI change in failure rate
A failure rate: failures per 1000 flight hours
A abort ratea
xiii
INTRODUCTION
Vibration has a recognized influence on the reliability and maintain-ability of helicopter airborne equipment. Airborne equipment failurerates - such as those associated with hydraulics, power train, structure,furnishings and flight controls - are expected to be related to the fre-quency, amplitude, and duration of the vibration environment. It is notreadily apparent, however, whether or not this effect of vibration ishighly significant or economically important.
The methods available for predicting reliability and maintainabilitystress the importance of the environmental effects which may degrade thereliability of airborne equipment. To date, the prediction handbooksprovide a constant multiplier factor K for ranges of environmental effectsto be applied to laboratory or bench failure rates; however, these K's
attempt to combine in one value the effects of humidity, altitude, shock,vibration, sand, dust, etc. Because different airborne equipments areaffected to different degrees by each environment, more accuracy in thereliability prediction can be attained by developing failure rates aroundeach environmental factor.
Component failure rates in fixed-wing application are demonstrauly lowerthan those of equivalent or similar componentz used in helicopters. Thisobservation is verified to some extent by comparing fixed-wing and heli-copter failure rate- appearing in the Bureau of Nave Weapons FailureRate Data Handbook.-
Figures 1 and 2 illustr-te the comparison for hydraulic actuators andselected electronic components. This trend suggests that lower vibrationlevels inherent in fixed-wing aircraft lead to better component reli-ability, but other major differences in the applications alo exist. Atthe same time, however, significant reductions in maintenance have beenreported on commercial S-61 model Delicopters when they were equippedwith vibration-reducing equipment.
The deleterious effect of vibration on reliability is not readily sep-arated from other environmental effects. Starting in 1970, however, aunique opportunity to relate helicopter vibration data to field reliabilitydata became possible with the installation of rotor-mounted vibrationabsorbers on USAF Ct{-3 helicopters. Recorded reliability data and measuredvibration data on these type aircraft before and after the installationof the absorber were available. The measured vibration data were acquiredfrom a test program conducted on three CH-3 helicopters at Sikorsky Air-craft. Reliability data were acquired on operational aircraft currentlyin the field from the USAF AFM4 66-1 maintenance data reporting system.
The study reported herein evaluated vibration levels and reliability andmaintainability records for CH-3 helicopters with and without the vibra-tion absorbers installed. Two CH-3 helicopter populations consisting of15 aircraft each were selected for this study. One aircraft group wasinitially placed into service with the absorber and the other group wasplaced into service without the absorber; each aircraft group had
1
1500-
1000
inI
C-
z0
wF-
_Jw 500-
FIXED WING
H E L I COP T ER
I I I
0 50 100 150
RELATIVE COMPL EXITY
Figture 1- Fixed-Wirg Electronic Reliability vs iotriry-Winp, Reliability.
35-
30-
C 25-X FIXED WING
U20-
0 i15-
F-
0
10- HELICOPTER
0 1 2 3 4
RELATIVE COMPLEXITY
Figure 2. Fixed-Wing Hydraulic Reliability vs Rotary-Wing HydraulicReliability.
approximately the same number of flight hours for the time period coveredby the study. The vibration data and reliability data were generated andrecorded prior to commencement of this study. Changes in CH-3 R/M levelsdue to changes in vibration levels are summarized in this report.
Ir
|I
METHODS MD RESULTS
AIRCRAFT POPULATION AND DATA SEPARATION
At the time the study commented there were 15 H.-3 helicopters which en-tered service equipped with the vibration absorber. Thus, a correspondinggroup of 15 H-3 helicopters were selected as a control group without theabsorber. Figures 3 and 4 illustrate the helicopter and the vibrationabsorber respectively.
The AFM 66-1 data received from 1.PAFB contain information on all H-3 air-craft in service, and the computer programs at Sikorsky Aircraft weremodified to allow the extraction of those data which would be applicableto the study. For the "without-vibration-absorber aircraft," data weretaken from 15 aircraft serial numbers selected from the last group ofaircraft delivered without the absorber. Only data prior to January 1970,the date of initial delivery of vibration absorber kits to the USAF, wereused. Information for the "with-vibration-absorber aircraft" was taken,by aircraft serial number, only from aircraft delivered with absorbersinstalled.
The 2light-hour totals for each aircraft were determined from the aircraftlogs or from flight times prcvided by field service reports. In conjunc-tion with the determination of total flight time, the time spent in per-forming various missions was also considered since reported reliabilityand maintainability data may be seusitive to the particular mission.Detailed mission profiles were obtained for the two fleets, and no signi-ficant differences were found in the way, purpose, and length of time theaircraft were being used. Therefore, relisbility and maintainability sen-sitivity to missions flown vas assumed to be equivalent for both groupsof aircraft.
The total number of flight :aours accumulated by each aircraft over the14-month period covered by the AFM 66-1 data along with aircraft locationsare provided in Table I. The percentage utilization for each mission ispresented in Table II.
The geographical locations of the sample groups of aircraft suggest thatthey may have been exposed to large differences in climatic conditions andthat this would impact uDon the reliability and maintainability data stud-ied. The aircraft without the absorber are located in geographical regionsranging from the Tropic Zone to the Temperate Zone (l14ON Lat to 520N Lat).The aircraft with the absorber are located in geographical regions rang-ing from the northern portion of the Tem~erate Zone to a region above theArctic Circle, the North Frigid Zone (61 N Lat to 76011 Lat). Climatolog-ical surveys were investigated and the aircraft without the absorber wereexposed to mean daily temperetures ranging from 40oF to 890F, mean annualsnowfall ranging from 10 inches to 23 inches, and mean wind speed rangingfrom 6 kt to 9 kt. The aircraft vith the absorber were exposed to meandaily temperatures ranging from 5OF to 42.6 0F, mean annual rainfall of5.5 inches to 17 inches, mean annual snowfall ranging from 10.5 inches to36 inches, and mean wind speeds ranging from 5 kt to 7 kt.
5
r- w -~ -- - -----. ---.------------- ~-- -~---~-~ -- ~. -~
A.C
w.I-
'4
2.
/
6
- i 4
4~
NN
-4
~1.
4I
7
U%
tTABLE I. SELECTED H-3 AIRCRAFT, LOCATIONS, AND TIMES
~Accumulated
Aircraft Flight DateSerial Hours Entered
SNumber (Utilization) Service Location
Without Absorber (3/68 - 4/69)66-13284 525 6/67 Eglin AFB, Fla.
-13285 530 7/67 Forbes AFB, Kan.-13286 557 7/67 Eglin AFB, Fla.
67-14705 520 12/67 Forbes AFB, Kan.-14707 595 6/68 Shaw AFB, S.C.-14711 684 3/68 Eglin AFB, Fla.-14713 725 3/68 Eglin AFB, Fla.-14714 559 4/68 Eglin AFB, Fla.-14715 116 4/68 Woodbridge, G.B.-14716 125 5/68 Woodbridge, G.B.
-14717 126 5/68 Woodbridge, G.B.-14719 425 7/68 Forbes AFB, Kan.-14720 476 8/68 Shaw AFB, S.C.-14723 18o 9/68 Clark AFB, P.I.-14724 85 10/68 Clark AFB, P.I.
Total Hr 6228
With Absorber (1/70 - 4/71)
69-5798 490 4/70 Thule AFB, Greenland-5789 510 4/70 Thule AFB, Greenland
-5800 511 4/70 Alaskan Air Com.-5801 550 4/70 Alaskan Air Com.-5802 445 5/70 Alaskan Air Com.
-5803 405 5/70 Alaskan Air Com.
-5804 500 6/70 Alaskan Air Com.-5805 495 6/70 Alaskan Air Com.
-5806 450 7/70 Alaskan Air Com.-5807 415 7/70 Alasken Air Com.-5808 335 8/70 Alaskan Air Com.-5809 345 9/70 Alaskan Air Coi.-5810 300 10/70 Alaskan Air Com.-5811 230 11/70 Alaskan Air Com.-5812 190 1/71 Alaskan Air Com.
Total Hr 6171
8
TABLE II. H-3 AIRCRAFT MISSION PROFILE
Mission Percentage Utilization
Training 33.6
Rescue 9.6
Logistics h8.8
Other 8.0
The climatological summaries for each location cited in Table I are con-tained in Appendix I. The ranges of climatological conditions cited aboveare within the specified range of values for the H-3 aircraft. The dataaccumulation period for each population extended over a period of 14months (more than a full year), so that it cannot be argued that theabsorber aircraft benefited by postponing maintenance from winter monthsto the more favorable summer months. It appears that the climatic differ-ences tend to favor the group of aircraft without the vibration absorber(warm/temperate) as opposed to the aircraft with the absorber (cold/arctic).In consequence then, it would appear that the values of reliability andmaintainability would be biased in favor of the aircraft without the vibra-tion absorber and may reduce the apparent effect of lower vibration levelson reliability and maintainability values for aircraft with the absorber.
The conclusion of this report as to the beneficial effect of vibrationreduction on improving reliability and maintainability may therefore beconservative when considering the climatic difference.
l, DETERMINATION OF VIBRATION MAGNITUDES
The measured vibration data were acquired from a test program conductedon three H-3 helicopters at Sikorsky, whereas the reliability data wereacquired on operational aircraft currently in the field from the USAFAFM 66-1 maintenance data reporting system. The question arises as tothe rigor of using vibration data taken from aircraft different from thosefrom which the reliability data are taken and pooling these data to formthe basis of this study. It has long been an established procedure inthe aircraft industry to acquire various data on a sample of aircraft andto apply the results of these data throughout the entire fleet of aircraft.
Vibrations are induced in the helicopter and its components by the mainrotor at a frequency = Hz or F n/2n
9
F nfl ()n
where Fn = frequency, radians per second
n = number of rotor blades
= rotor speed, radians per second
Vibration data used in this study were taken from the flight test resultsrecorded on H-3 helicopters both with and without a bifilar vibrationabsorber. These tests measured the vertical and lateral vibration ampli-tudes and directions at various points through the aircraft at the domin-ant frequency of 17 Hz , and values obtained are shown in Table III (asextracted from the test report).
TABLE III. AIRCRAFT VIBRATION RESPONSE*
Without Absorber With Absorber
Butt Water Vertical Lateral Vertical LateralStation Line Line Accel Accel R Accel Accel ax
(in.) (in.) (in.) (g) (g) (g67 () (g; (..
95 21(RT) 107 0.17 0.31 0.35 0.11 0.29 0.3195 21(LT) 107 0.20** 0.31 0.31 0.09 0.29 0.30
187 39(RT) 107 0.17 0.22 0.28 0.09** 0.16 0.16187 39(LT) 107 0.17 0.22 0.28 0.17 0 16 0.23243 39(RT) 107 0.22 0.32 0.39 0.05 0.19 0.20243 39(LT) 107 0.24 0.32 0.40 0.16 0.19 0.25290 39(RT) 107 0.19 0.33 0.38 0.35 0.13 0.14N90 39(LT) 107 0.13 0.33 0.35 0.16 0.13 0.21379 39(RT) 107 0.25 0.23 0.34 0.15 0.17 0._3379 39(LT) 107 0.17** 0.23 0.23 0.05** 0.17 0.17243 10(RT) 181.5 0.39 1.16 1.22 0.23 0.42 0.49243 10(LT) 181.5 0.77 1.16 1.39 0.35 0.42 0.55290 10(RT) 181.5 0.75 1.34 1.54 0.25 0.45 0.51290 l0(LT) 181.5 0,56 1.34 1.45 0.15 0.45 0.47542 0 16o 0.24** 0.90 0.90 0.13 0.57 0.58709.5 0 225 0.19 1.90 1.90 0.24 0.94 0.97
*Vibration Frequency of Five Per Rotor Revolution
** 90 or 270 °
Since the aircraft systems and corponents are exposed to both the verticaland lateral motions simultaneously and experience the resultant effect,the lateral and vertical vibration components were combined to provide asingle resultant vibration value. Longitudinal vibration is not includedbecause past flight surveys have shown it to be negligible. The procedureused to determine the vibration response magnitude is provided by the
10
equations below.
Given Y = A Sin (Fnt (2)+ (3)
z = B Sin F tn
where Y = lateral acceleration
A = lateral acceleration amplitude
Z = vertical acceleration
B = vertical acceleration amplitude
= relative phase angle
then the total response is
R = A2 s (I- cos 2F t cos 2¢ + sin 2F t sin 24)
+ B-/2(l - cos 2Fnt) (5)
The vectorial addition of the vertical and lateral components of vibrationresults in equation (5) above and is valid for any phase angle €. Thephase angles used in cal-uiating the values shown in Table III were takendirectly from the flight test data to the nearest 90 ° . Setting the timederivative of eouation (5) equal to zero and substituting 0, 900, 1800,or 2700 results in equations (6), (7) and (8):
R = (A2 + B2 ) @ 0 or 1JO° (6)max
lId = A for A> B @ 900 or 2700 COmax
IRI = B for B> A @ 900 or 270' (8)max
The absolute value of Id (the maximum one-half peak to peak value ofthe vibration level) is Wependent upon the relative phase angle 0.
The equations (6), (7), and (8) were used to determine the resultantvibration levels snown in Table III. The phase angles for most locationsare 00 and 1800 where equation (6) applies, and the locations where phaseangles are 900 or 270 are noted by an asterisk. In these cases, thelarger of the two magnitudes (lateral or vertical) represent the vector
11
sum where equations (7) and (8) apply.
The resultant vibration responses for the 16 pairs of vertical and lateralvibration components are given in Table III and are mapped schematicallyin Figure 5 and in Figure 6 for the H-3 helicopter without and with thevibration absorber respectively.
Sample Calculation
Case I:From equation (5),
00 or 1800
2 2 cosF tR (A + B) n (9)
max 0 = F (A2 +B cos F t (10)at n n
Therefore, F t = /2, 3u/2n
giving WRa = (A2 + B2 ) (6)ma
Given: Vibration Level at Station 95, Butt Line 21(RT), Water Line107
= 0.31g,00 Z 0.17.gO 0° (Table III, with-out absorber)
RL 0.359
Case II:
From Equation (5),
= 900 or 2700
R = {A2(cos2 Fnt) + B2(Sin2 Fnt)} (11)
o dR =0 (A2 (Cos2 F t) + B (Sin F t))ax t n(B A2 ) Cos F t Sin F t (12)
Therefore, F t = 0, n/2, n, 3n/2n
Giving: ax = A if A>B (7)
= B if B>A (8)
12
LT. BUTT LINE RT. LT. BUTT LINE RT.40 20 0 20 40 40 20 0 20 40
STA.50
COPILOT PILOT
0.31 0.35 100
150
D .28 .28C 200
O.40 .39c 250 1.390 01.22
: .35 .38C 300 IA5O 01.54
W L i181.5350
:*.23 .34I400
500
550 .90
WL 160
600
L i 650
(I.
* L =225750
107 WATER LINE 160 AND AB VE WATER LINEFLOOR LEVEL CEILING LEVEL
Figure 5. Total Vibration Response gWithout Absorber.
J.3
LT BUTT LINE RT. LT. BUTT LINE RT.4Q20 20 40 40 20 0 20 40
STA.
50
COPILOT PILOT0.30 0.31 100
150
3.23 .16 C 20
0.25 .20( 250 .550 0.49
D.21 .14C 300 .470 0.51W L 181.5
350
.17 .28 C 400
450
500
550 ().58
W L =160
600
650
700 .97
WL =225~750\
107 WATER LINE 160 AND ABOVE WATER LINE
FLOOR LEVEL CEILING LEVEL
Figure 6. Total Vibration Response gWith Absorber.
Given: Vibration Level at Station 95, Butt Line 21(LT1M Water Line
107
0.31L0 ° Y 0.20gL900 (Table III, without (13)absorber )
1 =41 o.31g (14)
The overall vibration characteristics throughout the H-3 aircraft weredefined from the 16 pairs of measured vibration data points by making alinear point-to-point interpolation or extrapolation (as the case demanded).
To illustrate the procedure, the interpolation and extrapolation of theceiling data points (water line 181.5) were carried out as follows for thewithout-absorber case:
At butt line 10 left the l.| values were assumed to vary linearly throughthe value of !.39g at station 243 and the value of 1.45g at station 290 toyield the interpolated value of l.hOg at station 250 and an extrapolatedvalue of l.46g at station 300. This is illustrated in Figure 7.
Similarly, at butt line 10 right the lax values of 1.22g at station 243and 1.54 at station 290 were assumed to vary linearly to yield an inter-polated value of 1.27g at station 250 and an extrapolated value of 1.60 atstation 300. This is illustrated in Figure 8.
Curves were then p ised through the station 250 data points to provide anestimate of the highest value near butt line zero and to fall off withconstant slope either side of butt line zero. This shaping of the curveis suggested by the data points obtained, which seem to show maximum am-plitudes at butt line zero near the rotor source of excitation. The lowervalues at BL-40 are also consistent with the lower measured levels on thefloor below, (Figure 5). This procedure yielded butt line zero values 1.35gat station 250 and 1.60g at station 300. This process is illustrated inFigure 9.
The peak value of 1.60g, station 300, butt line zero, w.-_r line 181.5 nearthe main rotor station, the center of vibration excitation, was taken asthe maximum value at the point, and the drop-off from 1.60g to l. 35g ingoing from station 300 to station 250 was assumed to continue at constantrate proceeding toward the nose of the aircraft. Proceeding from station300 toward the rear of the aircraft, a straight-line drop-off from the max-imum value of 1.60g toiR . value of 0.90g measured at station 5h2 wasassumed. In pr.ceeding Turther toward the rear, a straight line with in-creasing values ofrIAL was assumed until the measured value of 1.9g wasreached at station 75V5. This procedure is illustrated in Figure 10, andrepresents a logical way of connecting the limited number of data pointswith a continuous line.
Similar reasoning was used to establish the 1 L ax values along the butt line
15
1.8
E) INTERPOLATED OREXTRAPOLATED VALUES-
SMEASURED VALUES
1.6 -- I
w I
BUTT LINE iOL___
1.2350 300 250 200
STATION
Figure 7. Extrapolation, Interpolation Vibrationg Levels BL 1O(LT) WL 181.5.
16
IL
1.
INTERPOLATED OR0 EXTRAPOLATED VALUES
A MEASURED VALUES1.6,
* Il
1.e
1.4
BUTT LINE IOR
1.2 L350 300 250 200
STATION
vi
Figure 8. Extrapolation, Interpolation Vibrationg Levels, BL IO(RT) WL 181.5.
17
2.0-1~
STATION 300
1.6
1.21
STATION 250_____
.8 - - ~ ~ - -
80L 40L 0 4CR 8CR
BUTT LINE
Figure 9. Extrapolation, Interpolation Vibration g LevelsSta. 250 and Sta. 300, WL 181.5.
18
II
2.4 - BUTTI LINE 10
-WITHOUT ABSORBER
-Jw->w --
.8
0 200 400 600 800
STATION
Figure 10. Extrapolation, Interpolatior. Vibrationg Levels, BL 0, WL 181.5.
19
corresponding to 20 and 40 inches left and right of butt line zero at theceiling water line, for the floor level, and for the without and withvibration cases. Subsequent to the establishment of IRi values at them
floor and ceiling water line levels, a straight-line inerpolation wasused to establish values for the water lines between or beyond the flooror ceiling levels.
Figures 11 and 12 show the completed process for thr .ithout-absorber andwith-absorber cases, as described above, for the floor level and ceilinglevel vibration magnitudes.
Vibration profiles, Figures 13 and 14, were developed from Figures 11 and12 in order to portray the relative vibration amplitudes without absorberand with absorber at the ceiling and floor and provide a before/after pic-ture at these water lines. These illustrations portray relative magnitudeof the resultant vibration of the vertical and lateral vibration. Thedirection of vibration at each station is not indicated.
Ranges of stations, butt lines, or water lines are specified as the loca-tion for some of the components considered in the study. The linear as-sumptions as to the vibration magnitudes acting on the component wereapplied and evaluated so that the lowest and highest vibration level isshown over the range of locations for the particular component.
Sample Calculation Establishing F Level at a Point or Range of Points
Case I: Vibration Level at a point (without absorber)
Component: Nose Landing Gear Kneeling Control
Location: Station 270, Butt Line 20R, Water Line 190
Vibration Level: Station 270, Butt Line 20R, Water Line 107 = 0.38g
Vibration Level: Station 270, Butt Line 20R, Water Line 181.5 1.2 4 g
Change in g level = 0.86g, Change in Water Line = 74.5 inches
Ag/inch = 0.86/74.5 = 0.0115g/inch
Distance from Reference Water Line = 190 - 181.5 = 8.5 inches
Total g change from reference point = (8.5) (0.115) = 0.lOg
Vibration Level at Water Line 190 = 1.24g + 0.10g = 1.3hg
The same procedure is carried out for the with absorber condition.
Case II: Vibration Level at two points (without absorber)
Component: Anticollision Light
20
-F-
LT BUTT LINE RT STA LT BUTT LINE RT40 20 20 40 4Q 20 0 20 40
50
(31 0.31 .33 035 .35 100 .56 0.64 .64 0.61 .56
0.29 0.30 .30 0.30 031 150 .74 0.80 .86 0.78 0.70FLOOR
().28 0.28 .28 0.28 M.28 200 022 0100 01.08 0.96 .84WL =107 "CEILING
W L 181.5).40 0.40 ).40 039 039 250 )1.10 01.26 01.35 01.14 0.98
0.35 0.36 0.36 0.37 0,3 300 01.15 OIAI 1.60 01.37 01.08
$.27 0.29 .31 033 035 350 01.19 01.38 01.20
0.21 0.25 ).28 031 )34 400 0108 01.26 01. 9/4 450
.. 4 00.97 1.14 o .98
500 0-2 .92 92
5 5 0 4) 9 9. .90 .90
600
650
700 .0W L =225
750
Figure 11. Linear Extrapolation, Interpolation of Vibration g Levelfor Entire CH-3 Aircraft Without Vibration Absorber.
21
LT. BUTT LINE RT. STA. LT BUTT LINE R.T40 20 0 20 40 0 4020 1) 22.D40
50
30 0.30 . 0 031 .31 I00 15 0.36 .36 037 .39
0.27 0.26 .25 0.24 024 150 0.39 0.40 ).42 0.40 .37
).23 022 0.20 0.18 0.16 200 .42 0.44 0.47' 0.43 039CEILING W L0 = 11.5
0.25 0.24 0.22 0.21 0.20 250 )A6 0.52 (L52 0.46 .41FLOOR L = 107
.21 0.19 .17 0.16 ,.14 300 4O 043 49 0.47 ).40
0.18 0.18 0.19 0.20 .20 350 0.45 (.51 0.46
.17 0.18 ).20 e.22 023 400 0.45 ).52 0.
,J7 .18 M0 .22 .23 450 .7 )4.)..4 za-Q2"0 47 ()540o.48
500 0 5 6 ..56
550 * .58,
600 7/
6()0).71
650 .84
7009WL =225
750
Figure 12. Linear Extrapolation, Interpolation of Vibration g Levelfor Entire CH-3 Aircraft With Vibration Absorber.
22
700 650 600 550 50)0 450 400 350 300 250 200 150 100
WITRUIT BIFILAR ABSL41BEP
70 650 600 550 5M0 450 400 350 300250 200 150 100
WITH BIFILAP. ABVVRBER
Figure 13. Vibration Magnitude Prof'ile Without and With VibrationAbsorber at Cabin Ceiling Level, WL 181.5.
23
FWD
450 400 350 300 250 200 150 100
WITPUT BIFILAR ABSORBER
14150, 1100- 350 30 WD 25r, 200 150 100WiTH BIFILAR ABSORBER
Figure 14. Vibration Magnitude at Cabin/CockpitFloor Level, WL 107.
24
Locations: Stations 250 and 720, Butt Line 0 - 0, Water Lines 80and 225
Vibration Level: Station 250, Butt Line 0, Water Line 107 = 0.40
Vibration Level: Station 250, Butt Line 0, Water Line 181.5 = 1.30
Change in g level = 0.90
Ag/irch = 0.90/74.5 = 0.121
Distance from reference = (80 - 107) = -27 inches
Total g change = (-27) (0.0121) = -0.32g
Vibration Level at the point = 0.40 - 0.32 = 0.08g
The vibr-Ltion magnitude of the light at Station 720, Butt Line 0, WaterLine 181.5 is the exact measurement of g level made at the point (Table III).
Reliability Pnd Maintainability Data Source Used and Data Analysis
The reliability and maintainability data used to perform this study wereobtained from the U. S. Air Force Maintenance Management System3 andcontained the failure and maintenance data for the two groups of aircraftwhich are discussed in this report. These data were prepared and recordedby the U. S. Air Porce within the normal routine of aircraft operation andare considered complete to the extent required by USAF directives providedby Reference 3. These data were collected and recorded prior to the startof the study and were not specifically collected by the USAF in support ofthis study, nor were these data edited in any manner. Information pertain-ing to the aircraft discussed in the study has been extracted from thebulk data and listed in Table I covering a l4-month period of operation andrepresenting 6228 flight hours and 6111 flight hours for the nonabsorberand absorber equipped aircraft respectively. The reliability and main-tainability information contained in the AFM 66-1 tapes consist of date,job number, aircraft tail number, quantity of failures, action taken, whendiscovered, parts removed, how malfunctioned, man-hours, work performed onor off the aircraft, and work unit code number. Wr, unit codes identifypreventive maintenance tasks as well as components which required correc-tive maintenance. The data were sorted by work unit code, quantity offailures, and maintenance man-hours for each aircraft subsystem and com-ponent.
Because of the large number of discrete work unit codes assigned to theH-3 (approximately 2000), covering 37 general subsystem codes, the effectof vibration on reliability and maintainability on those items reflectingmore than 10 to 15 failures within each general subsystem code for the tensubsystem codes reflectin6 the highest number of failures or maintenanceman-hours is discussed.
The engine/powerplant subsystem was not considered in this study because
25
engine data is identified by engine serial number and is not traceable toa particular aircraft tail number.
The average failure rate for a given wo~k unit code was copputed by takingthe ratio between the total number of failures recorded and the total accu-mulated flight hours on each sample gro up of aircraft. Similarly, theaverage maintenance man-hour per 1000 flight hours, 14H/KF{I, for a givenwork unit code was computed by taking the ratio between the man-hoursrecorded and the total accumulated flight hours for each sample group ofaircraft.
The result of this procedure is shown in Table IV at the subsystem leveland in Table V for the highest ranked components within a subsystem.
Tables IV and V provide an overall view as to the dramatic impact of vibra-tion reduction at the subsystem level on reliability and maintainability.
The next procedure was to separate out from all systems, those classes ofcomponents considered to possess similar reliability characteristics.This was done for lights, switches, wires, plugs, connectors, hoses, lines,tubing, valves and relays. All UM 66-1 data were used and all items hav-ing failures recorded against them, regardless of the quantity of failures,were tabulated along with the computation of the average failure rate X andthe maintenance man-hours per flight hour, MMH/KFH. The intent of thisprocedure was to determine if a behavior pattern of vibration with respectto reliability and maintainability could be recognized other than thedramatic differences in failure rate and maintenance man-hours per flighthours evidenced in Tables IV and V. These results presented in Tables IV,VII, VIII, IX, X, XI had no readily discernible characteristic other thanthat which is evidenced at the subsystem level. The locations or rangeof location for all components can be related to the actual aircraft byreferring to the locating grid in Figure 15.
The term "average failure rate" was generated because it was not possibleto establish the type of failure distribution which fit the reliabilitydata. A failure distribution could not be established because time-to-failure information was not included in the AFM 66-1 data. Past studieson the reliability characteristics on major components of the H-3 heli-copter indicated an exponential distribution modified by early and wear-out failure phenomena. However, since a constant failure or an earlyfailure or wearout phenomenon could not be established for the data used,the term "average failure rate" is used.
26
TABLE IV. TOTAL AIRCRAFT SYSTEM COMPARISONRELIABILITY AND CORRECTIVE MAINTENANCE
Failure Rates (10-)I -MMH/KH
Aircraft W/Out With Failure W/Out WithSubsystem Absorber Absorber I Rate Absorber Absorber MMH/KFH
Airframe 223.7 107.8 115.9 592.3 209.7 382.6Drive 108.7 47.6 61.1 371.8 216.5 155.3Utilities 64.1 13.8 50.3 I06.4 26.3 80.1Landing Gear 91.5 44.8 46.7 289.6 189.8 99.8Lights 119.6 29.3 90.3 2h0.7 45.6 195.1Fuel 56.2 22.8 33.4 118.8 50.8 68.0Fit. Control 58.4 22.8 35.6 209.5 60.5 149.0Rotor 8o.4 51.0 29.4 321.4 278.8 42.6Coukpit/Fus. 33.1 9.9 23.2 48.9 23.2 25.7Electrical 35.6 12.4 23.2 79.4 26.2 53.2Hyd. Power 37.1 17.1 20.0 76.3 19.9 56.4Inter Comm. 39.5 21.2 18.3 71.2 49.7 21.5Radio Nay. 65.5 50.2 15.3 209.0 217.7 -8.7Air Cond/Heat 27.1 18.3 8.8 95.7 36.1 59.6Auto Pilot 28.4 16.6 11.8 94.2 88.6 5.6Emer. Equip 12.7 2.4 10.3 15.9 1.4 14.5Aux Power Unit 44.5 36.2 8.3 125.9 107.4 18.5HF Comm. 14.9 6.7 8.2 69.3 33.5 35.8UHF Comm. 23.1 17.6 5.5 67.9 93.1 -25.2IFF 8.2 2.9 5.3 21.9 12.3 9.6Misc. Comm. 8.7 4.7 4.0 13.4 9.3 4.1Weap. Del. 1.9 0.2 1.7 4.3 0.3 4.0Emer. Comm. 0.2 0.2 0 0.2 0.3 -0.1VHF 9.2 9.14 -0.2 38.8 36.4 2.4Radar Nav. 40.0 40.4 -0.4 163.7 188.2 -2h.5
i! * Minus sign indicates an increase in rate.
27
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The data samples were also sorted by aircraft tail number, and the totalaverage failure rate per aircraft and total maintenance man-hours perflight hour per aircraft were established. The purpose of this procedurewas to determine if any unusual differences in reliability or maintain-ability existed between each aircraft within a group as might be causedby variation in local maintenance policy, record keeping, extent of main-tenance facilities, or for any of numerous reasons. The aircraft totalsare given in Tables XII and XIII. The one apparent difference was indi-cated on two absorber-equipped aircraft located in Thule, Greenland (TailNumbers 69-5798 and 69-5799). The data indicated that there were no partremovals charged against these aircraft. However, the data on these air-craft were compared to the other aircraft in the group wiJth the partremoval data suppressed (on-aircraft actions only) on the remaining air-craft in an attempt to show equivalent types of data for each aircraftwithin the group. Total reliability and maintainability were compared onthe remaining 13 aircraft in the group and are also shown in Table XIII.
Using the values given in Tables XII and XIII, the mean values were cal-culated and are presented in Table XIV. Histograms were developed fromthe tabulated values in Tables XII and XIII and are presented in Figures35 and 36. From these histograms the 50 percentile and modal values weredeveloped as presented in Table XIV. These parameters, when compared for
the differences in the average failure rate and maintenance man-hours perflight hour between each aircraft group, indicate a change in A and MMH/KFHwhich is considered to be more than a coincidence resulting from the vari-ation cited earlier and could only be explained by the changed vibrationcharacteristic on the H-3 aircraft. Variations existing between likeaircraft at different locations under different geological and climaticconditions, etc., are fairly sizable, but are small compared to the dif-ferences between the means of the two fleets, when sample size is con-sidered. It is ccncluded, then that differenceb in R&M history are causedby the differences in vibration levels existing between the two groups ofaircraft.
The initial intent of the study was to consider the effect of vibrationreduction on reliability and corrective maintenance. The AFM 66-1 dataalso contained informatior related to the preventive maintenance performedon the aircraft. Since the data were readily available, the study wasexpanded to separate out 'he unit codes for preflight, postflight, periodicand special inspections; see Table XJ. The information displayed in TableXV shows an increase in the frequency and the maintenance man-hours forpreflight inspections and an increase in the maintenance man-hours expendedfor postflight inspection for the aircraft equipped with vibration absorber.This observation is not consistent with the previous discussion related tocorrective maintenance. However, these data represent only the "look"phase of an inspection and not the "fix" phase. That is, any correctivemaintenance is chargeable and charged to the work unit code of the com-
ponent or subsystem that has failed. Indeed, the corrective maintenancedata show that there are fewer failed items on the absorber-equipped air-craft, and therefore the additional inspections that are performed or theincreased manpower can not be the result of chronic problems with thegroup of aircraft where the additional inspections are performed as a form
62
UN 0 .--- 00 -t 00 t- t- Hi 0 It Hi t- t- 0) CU
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4TABLE XIV. COMPARISON OF AIRCRAFT A AND MMH/KFH BYPOPULATION GROUP - ON AIRCRAFT ACTION ONLY
Failures Per 1000 Hr MMH/1000 Flight HrWitnout F- w.hw ou7 wInAbsorber I Absorber A
Arithmetic Mean 1462 586 3910 166350 Percentile 1150 600 2750 1750Mode 1150 600 2750 1575
Arithmetic Mean: Equal to sum of the values for the individual aircraftdivided by the number of aircraft.
50 Percentile: Median Value - the middle value for the 15 aircraft.Mode: Value occurring most frequently.
i
65
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66
AVERAGE FAILURE RATE (10 - 3 ) AVERAGE MMH/KFH250 150 50 COMPONENT G 200 400 600
I I I I 1 NAME LEVEL* I I I I I --
I AIRFRAME 0.97/1.230.36/0.49
[ DRIVE 1.36/1.560.50/0.56
UTILITIES 0.29/0.48
I LANDING 0.41/0.32GEAR 0.23/0.28
r LIGHTS 0.42/1.180.27/0.63
FUE L 0.65/0.78
0.36/0.53
FLIGHT 0.79/I .32CONTROL 0.36/0.55
COCKPIT 0.34/0.73FUS. 0.25/0.49
0.51/ 1.14ELECTRICAL 0.34/0.40
HYD. 1.30/1.35POWER 0.49/0.56
INTER 0.33/0.75COMM. 0.22/0.52
RADIO 0.34/0.75
NAV. 0.30/0.51
AIR COND 0.74/1.30HEAT 0.36/0.52
_- WITHOUT ABSORBER
SWITH ABSORBER
* G Level given as Hi/LaRange
Figure 16. Comparison of Total Average Failure Rate andMWH/1F1 for Top 13 Aircraft Subsystems.
67
AVERAGE FAILURE RATE (0A) AVERAGE MMH/KFHCOMPONENT G
12 9 6 3 NAME LEVEL 20 40 60 80 100 300
ENGINE FWD. 0.95/ 1.11COWLING 0.43/0.4
ENGINE AFT. 1.22/1.42COWL 0.46/0.49
APU L33/1.63COWL 0.46/0.59
ENGINE 1.24/1.12 -WORK PLATFORM 0.43/0.48 I -
TRANSMISSION 1.19/1.321llllllll1 WORK PLATFORM 0A4/053
M.R. PYLON 1.35/167STRUCTURE 0.45/0.54-f!
ACCESS 1-33/1.63MU PANELS 0.45/0.57
SPONSON 0.12/035SKIN 0.10/021
M.R. PYLON 1.35/167FAIRINGS O.45/O54
AFT 0.25/071 _-RAMP 0.15/0.54
CENTRAL 0.35/0.93WS FRAMES 0.18/0AI
132 ....... 278.4
EI I I I l l l l l l l l l l A L L O T H E R 0 .9 7 L 2 3 A V G ) 2 7 8 .4
AIRFRAME 0.36/0.49(4V0 IIIIIIIlflUIIIIIIlUIBhIiUIU1EIII65.2 127.9
I- I~ WITHOUT ABSORBER
IDI~hIIUIl WITH ABSORBER
* G Level given as Hi/IoRange
Figure 17. Comparison of Average Failure Rate and MMII/KFHfor Selected Airframe Subsystem Components.
68
AVERAGE FAILURE RATE (10-3) AVERAGE MMHI/KFHO00VONENT G
10 8 6 4 2 NAME LEVEL* 20 40 60 80 100.M .. I I a I
IXMTR 0.561.59
CLUTCH 0.47
ROTOR 1.35BRAKE 0.49
LINES 1-9/0.58
XMTR 0A9
mOGB. FLEX 1.38/1.46HOSES 0.4W.55
M.G.B. I.4O/I.T6
SEALS 0.51/0.5659,8 AL128.1 1 7 -7
DRIE 0.50'0.56 IIIIIII32.4 YT
EZZZJWITHOUT ABSORBER
I1IUEEWITH ABSORBER*G Level given as Hi/Lo Range
Figure 18. Comparison of Average Failure Rate and MIIIfor Selected Drive Subsystem Components.
69
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79
AVERAGE FAILURE RATE(10 - 3 ) AVERAGE MMH/KFH
25 20 15 10 5 OCOMPONENT G*0 ip 20 30 40 50I NAME LEVEL
- EMERG. EXIT 0.95/1.02Now LIGHT 0. 38/0.5 6
POSITION 0.26LIGHT 0. 19
PANEL 0.44LIGHTS 0.35
FUSELAGE 0.05/1.9LIGHTS 0.08/0.97
ANTI-COLLISION 0.08/I .9LIGHTS 0.11/0.97
INT. GENERAL 0.75/1.02LIGHTS 0.34/0.56
CONTROLLABLE 0.29SEARCHLIGHT 0.09
LANDING 0.28LIGHTS 0.09
FLOOD 0.36/0.06LIGHTS 0.35/0.09
COCKPIT 0.69SPOTLIGHT 036
ALL OTHER 0.42/I 18LIGHTS 0.27/0.63
EZZJ WITHOUT ABSORBERi, WITH ABSORBER
* Level given as Hi/Lo Range
Figure 29. Comparison of Average Failure Rate and MMH/KFHfor all Internal and External Lights.
80
AVERAGE FA!LURE RATE (10-3
) AVERAGE MMH/KFH
5 4 3 2 I COMPONENT G* 2 4 6 8 to 12 14i I i i I NAME LEVEL I I I I I I ,
RAMP SWITCH 0.82
MGB PRESSURE 1.56 1.SWITCH 0.53
FUEL PRESSURE 1.48SWITCH 0.60
I MLG SPREAD 061SWITCH 0.35
MLG LOCK 028/0 37LIMIT SWITCH 0 18/0.20
1 BRAKE PRESSURE I 56SWITCH 0.53
RESCUE HOIST 0.68SWITCH 038
AUXILIARY SERVO 1.56PRESSURE SWITCH 0.53
'":XILIARY 0.42FUEL SWITCH 0.33
1AUXILIARY 0.41SWITCH 0 52
MLG UPLOCK 0 28/037LIMIT SWITCH 0.18/0.20
MLG DOWN 0.28/0.37LIMIT SWITCH 0 18/0.20
SHEAR TEST 074SWITCH 0 37
NOSE GEAR UP- 0.30LIMIT SWITCH 0.26
ALL OTHER O.74(AVE)SWITCHES O.37(AVE)
SWITHOUT ABSORBER
WITH ABSORBER
* G Level given as Hi/LoRange
Figure 30. Comparison of Average Failure Rate
and M4HiKFH for all Switches.
81
w 0 k
x ) 0 os-
j)
LO 0
w 0 ( 00
> CDCd
(D UI l n r) o i) w at) ) ~to 0f re C to C Dit)(J c C 0
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AVERAGE FAILURE RATE (10- 3 ) AVERAGE MMH/KFH20 16 12 8 4 COMPONENT G * 10 20 30 40 50I I I I I NAME LEVEL I I I
MG8 LUBE I 43LINE 0.52
IAFT RAMP "02" UPLOCK 05
UTILITY HYDFLEX HOSE o.3'UTILITY HYO 1.82TUBE 0.58
EXTINGUISHER 1.53/1.78LINE 0.57/0.66
MLG KNEELING 1.11HOSE 0.54
MLG ACTUA' NG 0.22/0.23LINE (A) 0.14/0.21
MGB FLEX 1.49HOSE 0.52
PRIMARY SERVO 1.56/1.72FLEX HOSE 0.57/0.56FUEL LINE0.37/1.35
FUELLINE0.36/0.52
PRIMMRY SERVO 1.56/1.72HYD LINE 0.57/0.56
ENGINE MISC 1.83LINES 0.54
v jAPO MISC 1.39LINES 0.49
MLG MISC 0.22/0.23LINES 0.14/0.21
MGB OIL COOLER 1.72HOSE 0.55
RESCUE HOIST 0.78TUBE 0.29
ALL OTHER 0.72(AVE)"=LINES 6 POSES 0.32(AVE)
" J WITHOUT ABS'o.tBER
WITH ABSORBER
* G Level given as Hi/LoRange
Figure 32. Comparison of Average Failure Rate andNMH/KFH for all Hoses and Lines.
83
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AVERAGE FAILURE RATE (10-3) AVERAGE MMH/KFH
5 4 3 2 I COMPONENT G 2 4 6 8 10' I NAME LEVEL I I I
VHF POWER 0.61 _
0 RELAY 0.34 0
ENGINE START 0.32RELAY 0.28
REVERSE 0.31 1CURRENT RELAY :.29
ENGINE ANTI-ICE 0 64CONTROL RELAY (A) 0.36
FUEL 0.61|FUEL RELAY
ARA-25 POWER 0.420 RELAY 0.31
f- DIMMER 0.61O RELAY 0.34 0
= GENERATOR 0.320 COPITACTOR RELAY 0.28 0
C ENGINE ANTI- 061ICE RELAY (B) 0.34
RESCUE HOIST 0.26RELAY 0.29 0
E Z1W!THOUT ABSORBER
I WITH ABSORBER
• G Level given as Hi/Lo
Range
Figir- 34. ri., of Av-re Failure RateArid I.CI/jFY1i for all "i lays.
85
-~-'~ -~ J
of insurance. The reasons which may explain the additional preventivemaintenance are left to conjecture; however, some reasons will be offered.
The average failure rate comparison for the highest retiked componentswithin each major subsystem as presented in Table V, are also presented asbar charts in Figures 16 - 24. These graphic presentations highlight thedramatic reduction in both average failure rate and maintenance man-hoursthat appears to result from the reduction in vibration.
INDIVIDUAL AIRCRAFT RELIABILITY AND M4AINTAINABILITY COMPARISON
The data contained in Tables XII and XIII were sorted in ascending orderof average failure rate and maintenance man-hours per 1000 flight hours.The individual aircraft reliability data were divided into subgroupsrepresenting incremental changes of 100 failures per 1000 flight hours,and the number of aircraft falling within each increment were countedresulting in the histogram shown in Figure 35. The same process was appliedto the individual aircraft maintenance data where incremental subgroups of500 IM4H/KFH were used; see Figure 36.
The histograms show a distinctinve difference in the average failure rateand MMH/KFH and suggest that the two populations of aircraft are madesignific-ntly different by installing the vibration absorber in the onegroup and not the other. A number of different statistical tests may beapplied to these data to verify the significance of the difference betweenthe means of the two (with and without absorber) populations. These testswill all confirm that the difference is highly significant. However,since there is not a universally accepted procedure for such a formal test,none is included in this report. These data also lead to the same resultat the aircraft level as are shown for the reduced vibration effect onreliability and maintainability at the component and subsystem levels.
An equally important observation which is made from Figures 35 and 36 isthat the improved reliability and maintainability characteristics shownfor the vibration-absorber-equipped aircraft could not be the sole result
of a difference in local maintenance policies and procedures. If suchwere the case, the reliability data for each group of aircraft would besomewhat the same since the inherent reliability characterictics for eachgroup of aircraft would be the same. The differences in reliability char-acteristics shown between the aircraft without and with vibration absorber,are, however, logically explained by a change in the inherent reliabilitycharacteristics resulting from a lower vibration environment.
LIFE-CYCLE COST MODEL AND LIFE-CYCLE COST DETERMINATION
The LCC model used in the comparative analysis is presented in Figure 37.Items branching off "life-cycle cost" permit evaluating in-the-pocketsavings, whereas those extending from "effectiveness" permit estimatingsavings due to increased individual aircraft utility. Improved utilityleads to LCC savings, since the aircraft productivity can be increased byimprovements in mission reliability and availability. Then, the sametotal fleet productivity can be maintained using less aircraft, i.e.,
86
5
LL< 4
WITHOUT- ABSORBER
3
0
zJ 2 n/
o0w /J
II1 o0
0-
4
LL
cWITH3r ABSORBER
tL0
zw
!0o 200 300 400 500 600 700 800 S 0 1100 1200 1300 1400 1500 1600 R"16 280t4700199 299 399 499 599 699 799 899 599 1099 1199 .. 29 1399 1499 1599 1699 1799 1899 23994799
FAILURES PER 1000 HOURS
Figure 35. Distribution of Individual Aircraft Failure RatesComparing Aircraft Groups Without and With Absorber.
87
n-3
01 WITHOUTABSORB ER,, 2
0
zw I
"0
6
-5
0
WITHIL
o ABSORBER
zW
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0 500 1OGO 15002C002500 3000 350040 4500 50055006r%10 65007000 7500 8000 8500)13500499 999 1499 199924992999 3499 399944994999 549959996493 69M97499 79,M8499 8999 13999
MAINTENANCE MAN-HOURS PER 1000 FLIGHT HOURS
Figure 36. Distribution of Individual Aircraft 1-2O1/FH ComparingAircraft Groups Without and With Absorber.
88
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increased missiop reliability and availability produces a savings in air-craft acquisition costs. Figure 37 also displays the data used in themodel for the H-3 nonbifilar aircraft. The LCC data, down to the thirdlevel of cost input (flyaway cost, GSE cost, etc.) represents 1971 valuesfor the CH-3E non bifilar configuration. Unscheduled maintenance man-hours and failure rates are taken from Table IV and adjusted to includethe engine subsystem. Scheduled maintenance values are taken from TabLeXV and adjusted to include engines.
Specifically, downtime is set equal to one-half of unscheduled maintenanceplus periodic scheduled maintenance, i.e., it is assumed that two men areused to perform these functions. In its present form Figure 37 shows thatdowntime is dependent on periodic scheduled maintenance which renders anaircraft unavailable for extended periods of time and is independent ofnonperiodic scheduled maintenance which can be done at the discretion ofthe flight crew so as not to interfere with flight operations.
The set of essential parameter values providing inputs for the LCC modelare documented in Table XVI.
The rationale for these life-cycle cost study assumptions shown in TableXVI is as follows. The CH-3E nonbifilar, 1971 cost profile was used asthe base cost profile in this study because it existed in the Sikorskylife-cycle cost data bank and was consistent with the aircraft type andtime frame in which the study was performed. The 10-year aircraft lifecycle is typically used by Sikorsky and throughout the aviation industryfor cost effective analyses. For example, under the Sikorsky contract forthe Army, DAAJ02-71-C-0046, "Expendable Main Rotor Blade Study," the Armydefined the aircraft utilization for the C/E study as 500 hours per yearfor 10 years. The aircraft utilization of 500 hours per year defined inTable XVI is representative of the H-3 family of helicopters. Indicativeof this are the H-3 bifilar aircraft stationed at Elmendorf AFB, Alaska.Between April 1970, and May 1971, these aircraft accumulated 5153.8 flighthours at a rate of 480 flight hours per aircraft per year. The missiontime of 1 hour 38 minutes, the aircraft payload capacity of 5000 pounds,and the mission cruise speed of 125 knots are those values specified forthe 200-nautical-mile mission for the USAF H-3 long-range helicopter. Theaverage operational payload of 50% is based on Sikorsky studies on CH-3Chelicopters opcrating under a spectrum of world temperature and altitudeconditions for typical cargo loads. The unscheduled maintenance value isdirectly taken from the nonbifilar operational data used in this study.It is the sum of column four in Table IV. Periodic and nonperiodicscheduled maintenance values shown reflect an average value of the non-bifilar and bifilar operational data. These values are derived fromcolumns four and eight of Table XV. The model was to consider that sched-uled maintenance man-hour requirements are the same for the nonbifilarand bifilar 11-3's modified by the difference between the scheduled main-tenance called for the battery absorber vs the bifilar absorber. Theaircraft failure rate is taken directly from the nonbifilar operationaldata used in this report. It is the sum of column one in Table IV. Allother information in Table XVI, except for the 24-hour operational day,was based on the nonbifilar and bifilar data printouts. Under hostile
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conditions, the 24-hour operational day would be typical.
The equations for mission reliability, capability, and availability usedin the model are as follows:
Mission Reliability = e (15)
where Aa = total aircraft mission abort rate per hour
t = mission time in hours
Capability = PL x V x OP x aircraft life x utilization (16)c ave
where PL = aircraft payload capability
V = cruise velocity in knotsc
OPav e = average operational payload in percent
1 OD - OD/DAvailability = OD (17)
where OD = operational day in hours
DH/D = aircraft downtime in terms of down hours per day
The following is a sample calculation using these equations with values
taken from Figure 36 and Tables XV and XVI. The values used are:
Aa = .0139 abort per hour
t = 98/60 hours (2 min of each 98 min flt is hover)
PL = 5000 pounds (2.5 tons)
V = 125 knotsc
OP ave = 50% (15)
OD = 24 hours
DqID = 2.675 dn hr 500 flt hr/yrflt hr 365 days/yr
= 3.6641 down hoars per day
IThe definition shcwn is consistent with the availability equation(22) cited on page 83 of Maintainability Principles and Practices, byBlanchard and Lowery, McGraw-Hill, 1969.
92
Aircraft Life = 10 years
Utilization = 530 flight hours per year
Therefore,-Aat (.0139 x 98
Mission Reliability = e a e () = .9776
Capability = 2.5 tons x .5 x 125 kt x 10 yr x 500 flt hi per yr
x (96/98)* = 765,306 ton-naut miles
24 operational hr per day - 3.664 dn hr Der day24 operational hr per day
= .8473
*Only 96 out of 98 minutes is used in forward flight transportation.
The reduction of helicopter vibration provides major savings in life-cyclecost. In the LCC comparative analysis of vibration sensitive subsystemsand componentq reported herein, the installation of the bifilar vibrationabsorber system saves approximately $350,000. This represents close to10 percent of the total aircraft LCC.
The cost savings statistics include 13 subsystems in order to encompassthe 10 most vibration sensitive subsystems from both the unscheduled main-tenance man-hour and failure rate standpoints. In each of the subsystems,the LCC impact of vibration reduction is shown for the categories of un-scheduled maintenance, spares, mission reliability, and availability. Thesavings within these categories are presented in Table XVII.
The total savings for the 13 subsystems is over $365,000. The investmentto realize this substantial LCC savings is an $11,000 initial cost for thebifilar system, and an estimated $4,000 operating expense for bifilarscheduled maintenance. Total aircraft LCC improvement may be as high as15 percent (one and one-half times the 10 percent), since the $350,000cited a is derived frm the 13 subsystems. These subsystems representonly two-thirds of the aircraft, and -tie reduction in evpenditures andfailure rates on the remaining subsystems suggests that appreciable ex-penditure improvement could be incurred. However, no life-cycle costanalyses were made to verify this observation.
A higher degree of confidence may be placed in the economies computed forunscheduled maintenance than in spares because they are based on well-documented 11-3 helicopter operational data. Lower confidence must beplaced in the spares expenditure reductions, since no direct data feed-back source was available indicating spares consumed. The spares costreduction due to the installation of the bifilar vibration absorber sys-tem was calculated by applying the percentage change in subsystem unsched-uled maintenance observ-d on the bifilar aircraft to an estimated Z ibsystem
93
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life-cycle spares cost. These subsystem spares costs are defined in TableXVIII. Confidence in the mission reliability and availability cost im-provements is of a level consistent with spares. In mission reliability,the lower confidence results from the relatively small number of abortfailures observed in the nonbifilar and bifilar operational data. Thenonbifilar abort failure rates for the 13 subsystems are also shown inTable XVIII. Changes in abort rates due to bifilar installation were com-puted by applying the proportionate changes in bifilar and nonbifilar sub-system failure rates to the abort rates of Table XVIII. In availability,the lower confidence in results stems from the uncertainty in the opera-tional demand and squadron size. Operational demands may vary somewhatwithout requiring change in fleet size (and change in aircraft acquisitionexpense). Availability is computed based upon the assumption that an air-craft will be down (unavailable) 1 hour for every 2 hours of unscheduledand periodic scheduled maintenance man-hours expended (on the average, 2men are used for maintenance). Equations for computing mission reliability
and availability were shown in the preceding paragraph.
Life-cycle savings per aircraft for select component types are shown inTable XIX. They are displayed in the same format as Table XVIII and theconfidence in the numbers similarly applied. For these component typesthe savings is about $180,000.
In the life-cycle cost savings listed earlier, these savings were based ona helicopter cost profile in which the development costs were shown as"written off." This is because the CH-3E is a production aircraft and thedevelopment cost (the cost incurred during the RDTE phase for developingthe total aircraft system) is a past cost that no longer impacts on air-craft buying/selling decisions. The bifilar installation as it impactsthe CH-3E is a design modification, and under the cost profile shown inFigure 38 may be included as an acquisition cost "delta.' This cost"delta" could be further categorized into nonrecurring costs and re-curring costs. For the bifilar installation the nonrecurring cost, whenprorated over 100 aircraft, is slightly in excess of $4,000 per aircraft.The recurring cost per aircraft installation is about $6,500. The costdelta for the acquisition of a bifilar CH-3E is therefore approximately$11,000.
95
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VIBRATION DESIGN AND ACCEPTAICE TEST SPECIFICATION ASSESSMENT
The objective of this study task was to assess applicable vibration designand acceptance test specifications in light of the reliability and vibra-tion data presented.
The procedure used to perform the assessment consisted of examining thevibration reqairements cited in each of the specifications applicable tocomponents and systems used on the H-3 aircraft. Twenty-nine documentswere reviewed and are listed in Table XX.. These documents are generallyapplicable to all helicopters and are not necessarily unique to H-3applicaticns. Three documents were singled out for detailed assessment(MIL.-H-8501(A), MIL-STD-810B, and MIL-STD-781B) because they specificallycite the vibration requirements for helicopters, methods for environmentaltests such as vibration, and tests for reliability where a vibrationenvironment is simulated. Excerpts from MIL-H-8501(A) and MIL-STD-801Bare contained in Appendix Ii. (The information on reliability testing inMIL-STD-781B is quite extensive; reference a complete copy to support theensuing discussion).
The specification tree provided in Figure 38 illustrates the interdepend-ency existing between various component vibration specifications as wellas the paragraphs within the specification where the requirement is cited.
Current specifications for the acceptance testing of a helicopter compon-ent's ability to endure vibration are inadequate. Components are testedaccording to fixed-wing oriented specifications by employing relativelyhigh frequency vibration (above 25 Hz). By contrast, helicopters arehigh-amplitude, low-frequency machines, the predominant frequency ofexcitation being between 10 Hz and 20 Hz. Further, helicopters are de-veloped according to specifications which address the maximum allowablevibration levels in the cockpit and personnel cabin areas only. Lastly,in some cases, inventory helicopters do not conform to the military spec-ification concerning vibration. A model specification is written declar-ing the higher vibration levels acceptable. Consequently, it is under-standable why helicopter components, most of which are installed outsidethe cockpit and personnel cabin areas, often suffer premature failures.Further, understanding of the influence of excitation frequency on fail-ure rate is needed before more specific recommendation can be made re-garding military specification changes.
98
TABLE XX. APPLICABLE SPECIFICATIONS WITHVIBRATIO, :EQUIRE?_NTQ
Suecificatior Title
M.LIL-H-8501(A)l Helicopter flying & Ground Handling QualitiesGeneral Requirements For
IfL-S-8698(ASG)(l) Structural Design Requirements, Helicopters
MIL-T-8679 Test Requirements, Ground, Helicopter
?.IIL-D-23222A(AS) remonstration Requirements for HelicoptersMIL-STD-8loB(4) Environmental Test MethodsMIL-W-5013H(l) Wheel and Brake Assemblies, AircraftMIL-B-8584C Brake System Wheel DesignMIL-T-50o1 F(l) Tires, Pneumatic, AircraftMIL-L-8552C(2) Landing Gear Shock AbsorbersMIL-F-18372(Aer) Flight Control Systems Design Installation &
TestMIL-C-18244A(Wep) Control & Stabilization Systems, Auto, Piloted
Aircraftz.IL-T-6396C'ASG) Tank, Fuel, OilMIL-T-5578C(2) Tank, Fuel, AircraftMIL-F-17874B Fuel System, AircraftMIL-I-18802A(Wep) Fuel & Oil Lines, Aircraft, Instellation ofI-IL-T-5955C Transmission Systems, VTOL-STJL, General
MIL-I-18373A(AS) Instruments & Nlavigation Equipment, Installationof
1IL-.H-544OE Hydraulic Systems, Aircraft, Type I & II,Design, Installation
MIL-H--8775C Hydraulic System Components, Aircraft & GuidedMissiles, General
MIL-C-5503C(3) Cylinders, Aeronautical, Hydraulic Actuating,General
MIL-E-708OB(3) Electrical Equipment, Aircraft, Selection &
InstallationMIL-E-25499C Electrical Systems, Aircraft, Design &
InstallationM.IL-L-6723B Lights, Aircraft, General Specification
MIL-STD-202D(1) Test Methods for Electronics and ElectricalComponents
MIL-E-540OM(2) Electronic Equipment, Airborne, General Spec.
MIL-I-8700A Installation & Test of Electronics, Aircraft,General
MIL-I-8677(Aer) Installation Armament Control SystemsMIL-H-18325B(Aer) Heating & Ventilation System,
14IL-STD-781B Reliability Tests: Exponential Distribution
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The impact of vibration on rtliability as an independent influence may causea failure rate which varies with time rather than one which is constant withtime. This is to say that reliability would be expressed as
t I I
Ao- (44 - )dt (8R =e 0 (18)
-At
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where R = reliabilityA = constant failure rate
X(t) = time varying failure ratet = time
In the reliability test methods of MIL-STD-781B, an attempt is made tonot only subject the equipment to vibratory stress at a specified level,2.2G ±10% peak, but also to include exposure time related to multiples ofthe specified MTBF. However, close examination of the test requirementshows that an equipment is exposed to the vibratory stress for approxi-mately 1/6 of the total test tirae (hot/cold cycling occurring at the sametime). Thus, MIL-STD-781B test methods also appear to be insufficient insimulating equivalent exposure time Lo vibratory stress since and equip-ment when airborne in a helicopter will be subjected to vibratory stress100% of the time.
Comparisons made between the vibration values cited in MIL-H-8501 (seeAppendix II) and MIL-STD-810B or MIL-STD-781B show a difference in valuewhich varies by an order of magnitude; i.e., 0.2g to 0.4g versus 2.Og to5.0g. It is apparent that general design and test procedures related tovibration are not necessarily designed with reliability in mind. Thediscrepancy noted in MIL-STD-.781B, relative to exposure time, shows up aweakness in the reliability test specification which requires correctiveaction.
A reading of all the specifications cited in Table XIX, and in particular,those paragraphs called out in Figure 38, will present observations sin.-ilar to those above with respect to vibratory stress levels and exposuretimes (other environmental influence notwithstanding).
The above observations suggest that components specifications shoullrelate more closely to the vibration environment that prevails in heli-copters. The above observations also suggest that the allowable vibrationspecification for the helicopter (MIL-H-8501) should deal with the wholeaircraft and not just the crew and passenger compartments. By having amore favorable vibration environment throughout the aircraft, componentR/M would be improved.
It is suggested that a happy meeting ground exists between making com-ponents more tolerant of vibration and making the whole helicopter lessof a vibrating machine.
105
MAd
DISCUSSION OF RESULTS
RELIABILITY
On the basis of the data utilized to conduct the stud. of vibration effects
on helicopter reliability, the reliability improves with the reduction ofvibratory stress.
In general, a component subjected to a very large number of vibratorystress cycles will require a small percentage improvement in strength orreduction in vibratory stress to to from a severely limited life to anacceptable period of useful life. Many components exhibited a large per-centage reduction or zero failures in the absorber-equipped aircraft ascompared to a significant number of f.ilures in the nonabsorber aircraft.
Table XXI compares the percentage chiange in the reliability to the per-centage change in the average vibration level for the 13 subsystems con-sidered in the study. The manner in which the one ratio is related to theother is unknown. The overall data presented in Tables V through XI, andFigures 16 through 34 suggest the response of reliability to changes invibratory stress will possess different slopes for different types ofcomponents, dependent upon their construction, material used, method ofinstallation, and location in the aircraft.
Overall changes in the entire aircraft reliability with respect to changesin vibration can be observed from Table IV. There was a 48% reduction inthe total aircraft failure rate between the H-3 helicopter without andwith the vibration absorber.
The evidence indicates, in all comparisons made, i.e., component level,subsystem level, and aircraft level, a decreasing failure rate with de-creasing vibratory stress level.
Several componerits listeA in Tables VI through XI did, however, show asomewhat higher failure rate for aircraft with the absorber. This effectis prevalent primarily in components located in the forward portion of theaircraft. Prior to adding the bifilar absorber to the 11-3, a batteryabsorber, momted in the forward equipment bay, was used to dampen vibra-tions in the cockpit area. (The battery absorber is standard equipment inthe without-absorber group of aircraft). The vibration amplitude in thenose o" 'he aircraft was nearly the same for both aircraft populations.The battery absorber is very effective in reducing the vibration in thenose of the aircraft but is ineffective throughout the rest of the aircraft.Because the bifilar absorber had a greater overall effect on reducingaircraft vibration, the battery absorber was removed from aircraft havingthe bifilar absorber.
Vibration data shown in Figure 11 are taken from aircraft equipped withthe battery absorber, and the vibration data shown in Figure 12 are takenfrom aircraft with the bifilar absorber installed and battery absorberremoved. (The battery absorber is so called because the aircraft battery,supported in special absorber mount, served to supply the mass needed to
106
TABLE XXI. RATIO CHANGE IN AVERAGE FAILURE RATE ANDRATIO CHANGE IN AVERAGE VIBRATION LEVEL
System 1 /
Airframe 0.52 0.56
Drive 0.56 0.35
Utilities 0.78 3.56
Landing Gear 0.51 0.63
Lights 0.65 0.53
Fuel 0,59 0.60
Flight Controls 0.52 0. 42
Cockpit Fuselage 0.70 0.55
Electrical 0.65 0.52
Hydraulic Power 0.54 0.54
Intercommunication o.46 o.66
Radio Navigation 0.25 0.75
Airconditioning/Heating o.44 -O,40
107
acquire proper tuning and damping factor).
The heater ignition unit of the absorber equipped aircraft also showed aninordinate increase in failure rate, and it is suggested that the increasewas caused by the requirement to use the heater more frequently in thenorthern latitudes where the absorber-equipped aircraft are located.
The basic objective of showing impact of vibratory stress on the relia-bility characteristic of airborne equipment has been met and the tabulatedand illustraed results strongly suggest that the reliability improvessignificantly when significant reductions in vibratory stress are achieved.
The data suggest that the useful life of an aircraft can be extended with-out the need to strengthen or redesign certain airborne components to with-stand vibratory stress if adequate methods of damping vibration or isolatingequipment from v" atory stress are used. This assumes, however, thatadequate testing .,ich simulates the actual vibration environment is alsoconducted.
MAINTAINABILITY
Corrective maintenance performed on an aircraft is a direct function ofthe reliability inherent in the design, and it follows that for any im-provement made in relia ility a proportionate reduction in maintenanceshould also be achieved. The data analyzed in this study show that in allbut a few cases drastic reductions in z.intenance were evidenced as adirect result of the reduced vibratory stress and the increased reliabilityresulting therefrom.
However, the reduction in average failure rate does not fully account forthe reduced maintenance because in addition to a lower frequency there isalso a lower mean-time-to-repair in some cases. For instance, using datafrom Table V the average maintenance man-hors jer failure expended againstthe central frames assuming the same size repair crew, is given by
M11H MH/!OOOFH (20)
Failure Failures/lO00FH
For the without absorber case then
MMH 83.2 -. 6.0 (21)Failure 5.0
and for the with absorber case then
- 6.1 = 2.3 (22)Failure 2.6
It appears that the extensiveness of the damage incurred (chargeable asa failure) is less in the with-absorber case than in the without-absorbercase, and thus less repair work is required. This also implies, thatalthough airborne components continue to fail or require corrective action,
108
the degree to which a component function is degraded and the extent towhich repair time is required to return the component to a functional sta-tus are also reduced due to the improved vibration environment. Thisobservation can be borne out by applying equation (20) to the data presentedinTables VI through XI.
Table XXII ;resents the ratio of change in 1,P.'/FH and the ratio of changein average vibration. The manner in which the one ratio is related to theother is unknown. The overall data presented in Table V through XI andFigures 16 through 34 suggest the response of mai.atenance to changas invibratory stress will possess different slopes for lifferent types ofcomponents, dependent upon th2ir configuration and location in the aircraft.
The study also considered the impact of the reductiou in vibration levelon the preventive maintenance tasks. There was an increase in frequencyand maintenance man-hours for preflight inspection and an increase inmaintenance man-hours for postfl-ght inspections in the absorber-equippedaircraft.
The reasons why this should occur in the face of reductions shown forcorrective maintenance are unknown, but either local preventive maintenancepolicies or a more rigorous operational schedule may account for this effect.Table XV shows that although there was an increase in frequency and main-tenance man-hours, the average number of maintenance man-hours per preflightinspection are 5.3 ,24H in the without-absorber case and 5.0 N24H in the with-absorber case. This implies a greatar aircraft operational frequencybecause prefligh+ inspections are done prior to the start of each flight.
The 50% increase in the maintenance man-hours for postflight inspection inthe absorber case is felt to be caused by local maintenance policy whichallows portions of the periodic inspection to be performed when a post-flight inspection is performed.5 This acccunts, in part, for the signif-
icant change in periodic inspection frequency and maintenance in the ab-sorber-equipped aircraft. However, it is also known that the USAF changedthe periodic inspection interval from 50 to 100 hours during the time per-iod coverel by the data, sc a large portion of the reduction in frequencyof periodic inspections may be due to this policy change.
It would appear that the ratio of the look-phase time to the fix-phase+ime would increas bpralise of the decrease in the ntnber of discovcrcdfailures; thus, preventive maintenance intervals should be guided by theamount of fixing required subsequent to an inspection. If, in the longterm, inspections do not turn up discrepancies, preventive maintenanceresources should be conserved by increasing the inspection interval.
109
TABLE XXII. RATIO CHANGE Ili -Mi/FH AN!D RATIO CHA11GEIN AVERAGE VIBRATION LEVEL
SYSTEM ... F w gi-!.r.1FHwl ° 9 W/o
Airframe .65 .56
Drive .42 .35
Utilities .75 .56
landing Cear .71 .63
Liehts .70 53
Fuel .57 .60
Flight Controls .54 42
Cockpit/Fusele 53 55
Electrical .67 .52
Hydraulic Power .74 .54
IC .30 .61
Radio Navj gation -. .32
Airconditioning/Heating .63 .40
110
SPECIFICATION ASSESSM IT
The specifications governing the requirements for designing to a stated vi-bration environment are suitable in terms of determining resonant responsesin an equipment, early fatigue failures, or gross deficiencies in mechan-ical design, but are lacking in requirements related directly to reliability.This is -,,ident even in the one specification written specifically forreliabili-ty testing, MIL-STD-781B.
A more deliberate and well-designed specification along with detailed pro-cedure is necessary such tha'u (1) helicopter vibration levels are speciiedthroughout the aircraft, (2) component specifications are related to theappropriate specified helicopter vibration levels, (3) the nature of thevibration level/endurance characteristics of different types of componentsare learned, a& d (M0 the statistical variability to be expacted among partsis taken into accouzit.
The change in reliability and maintainability under the influence of vi-bration can be dramatic, as is evidenced by the data presented earlier,.nd thus, it becomes important that design and testing for the vibrationenvironment be carefully scrutinized prior to planniug and performing reli-ability or maintainability demonstration tests.
Considerable work has been performed and documented relative to develop-ing vibration stress cycling curves for the many materials used in airborneaplications. However, except for special programs, little has been accom-plished in the reliability area for complete equipment by type, function,or degree of complexity. For example, the difference in the average fail-ure rates as related to difference in vibration level for the types ofcomponents shown in Tables VI through XI could not be generalized in a
series of curves of mathematical expression because of the lack of databetween the recorded points. However, this gap in the data could be filledby subjecting a large group of specimens to a test program which varies thevibratory strers over the range not covered. This kind of program wouldallow for acquiring the same basic data for components of varying types,functions, and ccplexities as has been done for base materials.
"111
ii
CONCLUSIONS
Comparison of system and component reliability behavior, as affected by areduction in vibratory stress, indicates improvements of 48.7% in reli-ability with a resultant reduction of 38.5% in maintenance due to a 54.3%reduction in vibration level.
The series of bar graphs presented in Figures 16 through 34 show thatthere are variations in the way in which reliability and maintainabilitychange with respect to change in vibratory stress even within families ofsimilar components. It is concluded that more discrete data, acquiredfrom a closely controlled test program, would be required on each groupof components to determine the precise characteristic of the variation ofreliability with respect to changes in vibratory stress.
The reduction in the frequency of failure with respect to the improvementin the vibration characteristic suggests that the useful life of aircraftcomponents can be extended without making any design changes in the equip-ment by reducing the vibration of the helicopter. This statement assumes,in large part, that failures caused by vibration are fatigue related,and according to Heywood4 , small changes in vibratory stress can result ina component's life characteristic being changed from a severely limitedlife to a reasonable life.
The improved reliability resulting from the reduced vibratory stressenvironment results in less corrective maintenance being expended on theCH-3 aircraft. This results in less downtime on the aircraft, therebyimproving availability and contributing to the reduction in the operatingcost of the aircraft.
The life-cycle cost analysis, based upon the data presented, shows thatLCC may be reduced by as much as 10% for the 13 aircraft subsystems con-sidered in the study, because of the improved reliability and maintain-ability brought about by diminishing vibratory stress. The reductionsare manifested by lessening the costs of direct maintenance manpower andspares, and by improving helicopter utilization.
Assessment of the various vibration design and test specifications revealthat they are inadequate relative to vibration design requirements andtest criteria that can be related to the predictiun of actual operationalreliability. The inadequacies result from insufficient knowledge of thevibration level/endurance characteristics of various types of aircraftcomponents and from the lack of vibration level requirements for helicopterzones other than the cockpit and personnel cabin areas.
112
RECOIENDATIONS
The following recommendations are made based upon the results of thisstudy:
1. Establish a vibration test program such that basic data can beacquired which will eventually allow the formulation cf thegeneral relationships between reliability and effect of vibicatorystress levels and accumulative cycles on helicopter borne components.
A seziple group of helicopter components such as lines, hoses, lights,primary and secondary structure could be used in this type of testprogram to expand the test to more complex or differently configuredcomponents.
2. Perform basic research and analyses in order to establish an adequatespecification which uniquely relates helicopter component reliabilityto vibratory stress exposure.
Ii3. Establish astudy/test pormto research and rptuonthe
various devices, both active and passive, which will reduce thevibratory environment of helicopter components.
4. Expand helicopter vibration specifications to cover all significantareas of the aircraft (not just those occupied by personnel) wherecomponent fatigue damage may result.
11!I
.1
113
LITERATURE CITED
1. Naval Fleet Missile Systems and Evaluation Group, Failure Rate DataHandbook, Corona, California, September 1, 1969, Vol. lB.
2. Paul, William F., DEVELOPM.ENT AND EVALUATION OF THE MAIN ROTORBIFILAR ABSORBER, Sikorsky Aircraft Division of United AircraftCorporation, 25th Annual National Forum Proceedings, AHS, May 1969.
3. Department of the Air Force, MAIIITENANCE MANAGEMENT MANUAL AFM 66-1,Headquarters USAF, Washington, D. C., 10 February 1970.
4. Heywood, R. B., DESIGNING AGAINST FATIGUE OF IrTALS, Reinhold,New York, 1962.
5. T.O. 1H-3(c)-6WC-IPRPO, USAF Series CH-3C, CH-3E, HH-3E Helicopters,Preflight, Postflight Inspection Work Cards AFLC, Robins Air ForceBase, Georgia.
- I
APPEN~DIX ICLIMATIC BRIEFS
TAEVL E. XI. W CLIMATIC' BRIEF,
LGILIN AFB, FLA.AS CL C8 T 6R/EF - SOi.C
R paeed be ETAC ( A, 1132 ELEMTO04 aj ft S7..,TR L Xa-7.F
TEMPE6RATLRE?F) PRCPSAT0c4 641'j MNO (i'T' FVE d MEAN NN45ER COF 0AIS
22 I -V 2- I
FES 1-A 4. 12t 3.7--.-[2 0 *'... 1 -; - e - -T -1 0,_ _ _ _
&JAN 0 .2 .1 0 0 ' 7 3 17'-'Ac )Ca 0, 6 2 25 0
. '6 3.23.1-23 J7IZO Gj1! 0
-. '71 oss'~;1~ a~ O-o457 15 ifi '04-.9 * 9-6i
OCT.3 95 58_t595.%M .3.350 L-551D-T0 10 ' 2 9 .. 1: I 0 0 6
WNe g4 7 3.6 3.2 0 a S~ 5 13156.. .35 350 72 0 2 12 0 22 0 5
9 V.93d uS. ')A I W7Am~ 1.. -/T iai e 6 . i; 7 -1 : r ~3 5 ~..z '.: b fe 0 *9 13-, W 171 2-. 5
EYR 12. -'.1. 2 .. 1.I. ' i It 52! t 25 ;. "L .10 1, 1 -I ' L
r-S.* Ra: Bowly~ bs - a=s 1.- 1eov S.
JIM! AAVj ICT AVSLJ.BS. -tISS T)1I. C~~AY, . .:St.4 cc,. '(~ ~P PPIA3z
RLYtNG *LATHEP (%FPEO; HOCMA ILS1) WP FEB WMA APR U -UN AUG SEP OCT NOV CEC AMa ErR
Z3 -02 _ 5"3'015 5 0'?1loss tk AA ...
0 3..O --.~ 39j7 1. -T 2' V1, zi-n - U* 1 9'
3 300 " r-2 11 %A 1' _0 - 2 ._ 21 _" !17 v; 2H 10
AM or 111 251 *'. n -.5 1. * I 31
l=nos 2 , ~r-,-21r 2) _ 3, - . - --- D
ZL 10 '0~1 25 *. 2 uY5
less thm A
2 - 1' 5 3. 3, Zl 25
-C- z Z5121.23. r 1 3 4 121- 1.
TSU
l~sstae ... 1.1~XIA~ 2 1--11 x10 1 a.
CIO 21 .1..Z 4 ?jJ8 4 1.4. 3 J i"2
201 IC 155 I 01 2
2CO foot LJJ.-
loss u n., 17 -6 d2..0. i 1
AV$ 62v sA LSP I$$-4.V0L...L...
115
AWLIMATARE I eq AM .
Prerd by ETAC 110 v.9S7 S4 ELEVATION. l14fISTtlLTRS U7CTEMPERATUREMF PRECIPITATrON(m)2 WkINO CXTI MEAN MEAN NU'X.8ER OF DAYS
JAM 69 37.,19 -5 712 N 5 I7 i9 6512CI.101700 611 . 2, fl 101 0
02TL 2 5
FED T?1 1-3' 1. 6 x 964 iSit 4-.2 118O0 8 11 31 i 1 1 1 ~ I2MAR 901 514 31- 6 2.82.6g N 353 3 .S8o 12 1 ~~ I .1 '6
JUN 9 yd11 .. 1. Ild 50 f65B3.26, 90,1 0 1 01 f -I,22 0APRt 956~.j33.~6 01 ~ 1120 9~ 231 , 5 f1iFt0AY 97 214.13.L Aq 15j~ oo n 3' 81 31 #1 0 6
JUI.L 111 8: 69 53 3.7TI..6 0.0o 5s VIt' z 36&61 92 01 0 81 4t13j9 01 0 59UG 1t 67' 183.5 3.3 0' 0 jS q '.0p 52 , a .6 0 13 50] 7 -2 06 5 115 iz7 3;
x.r lad 0 58' 33 Z.8V2.V, #1 fj3 42 60 1 -50" 14501 7 12[o 6 46i26 01 oOCT ~ ~ ~ _ T6j 1>5q 21.5'..1i#s T ILt. 2 ,oF6 2_f0 (P2, 6'i-623
Id 79t . 31. .2.3,Z 1 5 jS916'5~1.d175 6i i 0 01 1DE 6-1- CTO -!-- 13:0 ,3 i _9 i5
411 23 -9F 1. . .1W 3 62;ii7Ot '1AW 11- 5 4.5 -15132-4 4.95$23 2'S 91 %S; 7956 -3 16!3 ?3 22 15 5. -.7 4 s51 12.:1 5
JEYR 2ZD 220 2A2320116 620 1.2)- C O 21 2a ZZ 2: 20 . 6- k; lo it. - -. ;1
[.5.L MS= ~ 3l Ott: On~ 12 - Amm 17, Zec '.7 - -,-y %9, Ft- 52 - )br 65* bL1y Cbs: OctC,- 1= -c - 70-.4. ?:ec -0-ct i.
IITE, -jar &.AzL.5t Z. .Izs . .- .. *." !"C. ' 9.5 vV"N!T --t CIAlgE
FLYING OEATHER C%FREQ) mOURS (LST; .A.FE3 IMZ., .4;R MAY ,--s .U- AU., SEP ce U.A , E-
less 27 : 1 1. _ I .15 1 I___ -n---3_ - -- _Z 1_2 - a- A l_ ;Z_-,
3 --- 15 -17-- ZL _Z~ _ _1% 15 -A. 27
All z-- 16 I 1 1
legs - 3 .. 6 1..p3lla -23m 22 21 -'Cr-j:i17 1 . I S 2 62 22J
10feet .. 2 z~2 21-LZ 14_j . _t_ 7 LUX : 15./e 19 2:12 2 -- U _1. '' -
7:2 j.ae /~A or8 u_27T I
lesst!z 147 1 - 32j 2jLI ":3 17 1 u 5 2 4 7 V 2-11 1*- 13 1
lede than V 1 7. J 4 I *o f; i t
-.- E . - _ .-- 9 9 120feet C9._U9 u 2 1_4
LLVI .3-~ 16e~sod / ~ or. 14 Il 2.2
2~~~~1 oil" _.? 2C3 14 1L10 3
u62
TABLE XXV. AWS CLIAIC BRIEF, PARKS AFB, P.I .
AsClmATIMCOR/EF - ' m NI L- mimp. PH VLZ RI0O0 191.9- 63 WSN 984
iFopwlod by ETAC ( OCT 190 )K 1. 31 E 121 01 _______ ELEVATION: 8#. USTNLTRS RPM
TEMPOU.2UREMF PRECIPITATIONCIn) WINO (KT) MEAN MEAN NUMBER OcDAYSIT f -.. o TI~ :rEM RTUREK1~
Q. 0 0 A . .4:1 Itr-E
i Pi if.
3 .. l 7# W ili. x z__ _- _ _.-- -I - --
9_ . 7_ I _.1 _o i i _I ..
96 18 5 0.0.71 1 7 7 1 U 2,0 0 331 9 Y
F 90E 8 71 25 31-f lIf '30TAPR 5891.1 6b E 33 . ,6 1 2 071 62 0. . 3 1
MAY 9991. 17o 733 8 j .82/2 2 2 967 0/1 30 10 2
JUN 10191175 68 10.1p .91 1/ W/ 133 9367,7.f5 16 5 1 1813 021 11 . 7
AUG 948875168 17.019.0 w 27 9 (-.$TV 400 #7 M0 9
SCT 9--88 73 1 12 7i 0i 1 40 2 13 2 110 36 -,'
1 '1) .1z -2 l 3 r, 7 1 91 7.8 1 Wi--22 #.3 109.__t
OCT 1 i 7!b0± j_, . - 1 71 35 I 0t 715 13 31 O 7
NN z "-i-r-o -o+-1'-T"5 ", ,! .i -F t - ; & 9
DEC 921 C6 169 61 2.2J 4.h; E 4127 j3 65"flX .7)r 30 ...I 112.1 5107
ANN 0 7 O8 3.._j- Z 6 -40-91 o 672 -.740 37 i7 8 3i63 , -1.
-YR0 2 C 10 19 , o 15 5 1 ; 1 1 " ' 5
IIASJan Fab tA?Apr N*y ;u Jul Aug SeP Oct NOT DoC AMi
mbrObsorv doWtIin: 1/8 A/3 AI- A/3 A/ A- A P A/ A B A/B A/B A/ Ar
(PO8: 1945-1969) 60 i 0/0 0/0 p/') 1/1 2/1 2/2 2/2 1/0 2/0 5/1 2/1 2 1719
') 0haowsropial Stors 120 0/0 0/0 /.' 2/2 2/1 2/2 412 4A 2/0 9/ 9/ 6/6 40/28
/ o cly 24 -w. 1/o 0/0 1/1 4.-4 7/3 -/5 12/6 13/7, 11/6 18/13 22/16 11/9 10 7 70
RUSSM FOR: !Hrl7 Oh.: Jan5f-1c 63 ,sMLly Ob: !'Ar -9-Zwo 63 UJ'A wind) refers to highest Mnd speed interval-
FOTE; =: ATA PCI AVAILABLE. ILES-, TkIU .1.5 [.A-, 1.5 I.oW - -. OR C. NT (1 AS A -' - -'
FLYINGWEATHER (%FREQ) HOURS (IST) ijAN PEB MAR APR MAY JUN JUL AUG.SEP 'OCTI NGV7 DEC ANN EYR
loss tu0 ._ __ _
and or4_3 1 - 1 Z5 3 1 91 1 6
mM/a "1 1 7~ 5 1 3 1 317 1 7 1 13 _
less than P; 1o 7~ -t -- 2~- 9r 1 0141;13il1 - 2 F-o1 o',, 5- 8" T
lh HOGn S 2 2' 4 6' 61 9 6- 5 5 3
less then01500 foet1 0 /T- 2 )1 3 21 e
and / or 2
0k- 2 1 1.1 __ _ _
1000 foo 0___ - t2less /br 19-I3 1j 0,b -1,4 .j2
Is. OT o 0 0 , 1, o 1 4i0 1i 01lomthen 93 M-0 q, A U 0~F o
20 foot 063 - 6e ZSL 11 L 1 _O I2 1And / or - 13X of O -0 11 2 1 1 1
lets then 1 _ - I 9 7S 3L f. Art~ O 2 1 L
I AL WM 0 d, lil, 2
&' / 1 -1-- o -a d01 U ul
TABLE XXVI. AWS CLIMATIC BRIEF, SHAW AFB, S. CAROLTIIA
/MA7;ZC8R1EF !n PERIOD 192-66 8,8,? 4 9
Ptsomedby 0T*A(~. ns Jicno)11113 9 2 ELEVATION 262 tIISTN L7RS E=S,TI PErATURE,.F) PRECIPITATIO(m 1IND (KT) MEAN M NEANNUYEqOFAYS T
- III~ w, I ~ I ~ .,- "TEMPRAURE(*F1 c.K ! ... -:; i I , - -' 0 o MAXINU
JAN 81.57 36 10 3.0;J.8:# 1 :"J 6 /. 79'561361.217W. 91 Z 'Od 1Z i iO q 910O 6
FES 82 ZD 9 . ,,, I 3'J165 4 3 7t* 22j80) 0 9lil, 'i i lo i
M (895 0 .:. i...1t=s So f6 1/02/12 7l5 W3~ 7/0r 1 3 3 2,-bA.M 92.5--o,/2; L WIO3 17512096769.30760 7 9jj 07/612.120 i l'18
JJL 3 7a 59 90~5' 80 LtJAN 8957 33 62 3.00A.8d 1 6148 791/0610.2 70, 2f 0 0 4 Of 6#0 92 0 10 17ES 8i k 0 9 3.6. 3~ 650765 t371/.220150 2 tI A 1) 7
4"S. €9 rv. .3:s .3 V; 4 ., _s7 53 .7715014 :.26:5 1-, ."3 1 ,1,. 1J 2 ,1. 01 3- 310APR 99 7T 1 541 32ii 0 Ord 5 JU9. 3 O1 T v 1 og N -JUN 1061. 651 4.1-3 v 6.&P j86,s521.6&.553 9c 3 0 -0j 17 641 15 28 0 0 5WL 2 3 5 23 .213. '210. 0 r 25 25.955 25 2i 1 2 4 22 0 20 ' 131 23 33 23 a b
Jz...b , o - w, u ~ Set Oc'- T - t To .fle -A
AUG 02,8 71, rn 56 tjz 5.1331 A0 Av '53i A/ 57. /9 !A00 A/S A/ 301/s aaP 0 8: 619- 9 4 69 57'64 6/0 2/0 3/ 1 2/ 0 /0 ; , 0
OC 00 997 5 0i t,513_660 6 2 0L .,4
NOVi" 0 -2 W 166 451 6 1205 1 / 69 3. 9 5 1: 6 02/1. 1/0 G- 2 12DE.rCe0 1/7 1 0 / 10/8 9a1 36/16 1/ 7 1/ 1/1 1 11:
A G N02 L 4. 2155,1 : 1 I.i 1 , 3 0 17 21 1 1 .2 2 5-23123 123 !, : L 182589 5 2, 2 1 , 2- 20 20 2 '2 2
Jan /* 1or AprJ May29 Jt6 Jul 14 Se.2 t NOT)52 Dee iiilA)HuricntsTroic- Stm 2 .:z 2'8 1K/0 6/ 6 7 9/ 26/ 13 1/0"13 1L9 13
_____________C=17 2 lV0 2%N 2 10 10/ 10/. 2Z1/21 316 /7 10 21 1--2/41
N=F' .T The -VILU $LESS LY.1 0 . f 5 0. Pe I ;rrs R 0. PEjj W4S CJ7LLOR -- 4y U -.--..---
CIG 1 00~-0 0 47 22 21 .'91 s is 2,z
100 fet 221197 1 2 2 21_' ; 7_) z.
§7= H- t-21-'173 1 -!. LA , 3 1 25 2 1) 73 2? -- - 81
A___ _IL 0U 21 :2 1720150 1 4 15 91912 IS 13 1 300029 81L31.7 7~j6 6 13 1 L 13 10 1
Cieethan 93--03 7j... 1 21 1212 1k~05 i 7 i1000 f~et -02- 1 i~ A jA B 2 0 S 9
and / or -J [-~_6 : j :1jKV6L10-I11 5 , 1 2 f 2 3. A
lessl tu 15j, - 61j 4 3 6 71:-7AlLo S 20 , 13 1361 . 6 4 5 7 -
CIG -iZ-32-A 2 7 ' T1I 61 -st ;n 0 n 3 1
lse ha a/or -1 __t a-_.- 1- fT 4_L
1000 ~ ~ ~ ~ _ fot 1 0-9 2 7 1I 4-1 91-1 -
a od /'.' 63 09-11 5O-I OL.IL..........10
12 2 21 .2--'~--
TABLE XXVII. A'S CLIMATIC BRIEF, ThULE AFB, GREENIAND
WSCL.MATICBRIEF "" P'0RIOD1951-65 JWBAN,, ~ ~ ~ ~ ~ ~ M *,,, ++o) 04+202Prei d by ETAC (OC 1970 s 2 68 &s ELEVATION: 261 fI STN LTRS
EATUREW PRECIPITATION ) W N T.; MEAN I MEAN NUMBER OF DAYS
hi.~~ ~ ~~ Lt 0 o. 'i-- . U1iUdIIMW
7ZIIXZ~ o !E. Z;:15 w .2 0
_ AN .1171, 9 0.5 11.7. 51 15 1 80 8160 1 1 # 1 5 - 0 1 2_. . 5 A ' -
MA 3 L - . 2 .0 . 21 1 5?-_Ol1,5_ 0 _ 0 6
APR _1 8t_ 1. 0.2 !0.3t 2 Sh 63 43. .2,13-0 4 0' )' 0I 4_ f- 3qAMAY 27 28 171-11 0-0-_8__ O50_JUN 63 14 321 22 ia 0 . . 1. 3-w 6- 71284~4.5 1 0 0 11 4 17 o 6JUL 6216 3 26o.8,2.o' 1 _w 6 64 1-91_1000 .5- 1 0 2_ 8_ 6
A62 362 fi1. t 2 W OD 7 0kl - F 90
1E I9 321 23 _.. O613 7 Z 6 727 .09 )01 0 8 -6 I 0 D 26
L~C 3~.15-33 ~ 7 -6766.i.O25O OL72 .0 -: -I z -IQ,
ANN- 630 1917 -"hk 0.. 6 6 ' 1 51 T7~o6 6~O 10 5 I 2 !o k ,1.017
o o3
Htza Ocs: Noq 9
NOV '. ,,I0 51..., Z 7-L)22
les .bsn .6 1.. 3 2: 2. O1~1 ~79 1
__ -- - -Io.____ _5,.__9_ 301 0:; _ca . Ia rq- ,-2' 5.15 =3 9. 7"70 50_, U a
AN 6 3 5 ,E8 2 6 10 .62.1 1 001170 1 6'S ' : , , 2 1 3' 1 5 1 1 9 11 1
AJS,.WO FOR: Hourly Cbs: N ov 51-Oct c5
D&O1y Oby- -N 51..14a? 8 L.,rUOTE; -,%. TA tUOt AVAILABLI . $LESS i1 0.$ DAY. 0.- no 0.05 ili-OP 0.$ ,;_Rr_,1T Wt) AS A-IPLICk+-LE.
FLYING WEATHEWR (0/c FREW) HOURS (,S T),jAN FE9 -MAR APR MAY JUN JULL AUG SEP O3CT Nov, OEC ANN EYR:
CIO . L,__ I t8 .._ ' . 9_ 23 2 2 . 1 * lb . .2 ,-5o '
less than 03l _ .7 812 j..J 7_6_17 . 1
r9- 21l 01L-9 --. 11500 feet 2__4L - 9__++_, ,, 3,1, 11 71wA/or 4-22 _L .133 2-. . 55-_1 0_.* - -. 6_ 1,
+
Ve,, ,~ i-.._5-1_LL3.l I_..LO . .. 2L. 6._ 1 - _. 7.__ . 1011q _ 53 'a J
' st .182. . 11.121 a* a_,.l.b i + 16 126 89 10B X .. 6 t "_.
3 UI.. 21 15k. 6 9 10' . 11
leg~~ta IS-17 1 04 6.j0 21 2 3. 5 I 6 IQ 91_~_% 2
_11 .. _. --. _ 17 17_ 1 ____. 11000 fee- . _ 7 -_AL _. -51k' 2 1 . IS - IQz- I _9 1
1 16 8 8 2 5 2 015 18 S 16 1-
less t! a18-. _ 22.12 5_6 9 1 4 1U 2412 + 6 " 10 . 6 T0 ; "
2u o .... ... Li - "2
.... 9 .. 3. -k. 10 - . 4 . . _- 6 7- 8j""
le .21 -X5= 7 2 13 7-5 10 16 16 " ' 5 O 1 6 5 8tl kI _ f41 61 -S 2 1_2- ' 91 6+I .d 1 * 6 L IALL HiOURiS3 11& .6 1Jo J.713. 219 1 3 1 .9-to 1 2 t..-
---d/or -021. _6_ 2 .. , _1. L _ ,.
-5W6-X-! _5 6su 9. ~ 2 3.k l
Ae, a±1e I..._,,. .L ... 4. O-.3 4-... _.+- - L.4-. L t L -L21-____Io a I _6I 7, 2; 2 -1 I Is12 1- 2,h z 3 1k
AWl -e:, lt AWSP lOS-c VOL .. L..-
var __8 1 ~1191-&-
TABLE XXVIII. AWS CLIMATIC BRIEF, WOODBRT:'Th AERODROME, EUGLANDWSCLIMATCBRIEF , (j.n-m, vm PERIOD 195.-70B WAN *35051
Pro d by ETAC( JAN 12 IN 52 05 E 01 21 " Lo ELEVATION: 95 STN LTRS ZOVOITEMPERATUREOF) PRECIPITATION(in) WID (KT) MEAN _ MEAN NUMBER OF DAYS
JAN61 8 16 I . 1 190TEM 1TURE(F
1. . '"I2 87 1JUN +J 3 1.70. 0 ;0 So75o1219.570 91 O0220I007
JU L !sZ - 2 w.1.7o!o8'
JAUG5L 3 _j4 21p84149 4 8OJ 7 1831 6'150 16ii1h O13 9 01 o o 13 v1592 .17 0k0 1!J~ 1535083 lizO 02 11101 0112 007
FOS 5 3 1 5 1..I 3 18162 W7 3.215 6LMM '36 4816 .2CL.1 1 31J. 9 150 35 .20U110 161 Ii _f____15101_1_9_0DEC 2 0j7 .. . I I .S 8151 5 39 -210 13 J 1 0 410 031:0 7
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APPENDIX II
EXCERPTS FROM ASSESSED SPECIFICATIONS
MIL-H-8501A(l) Helicopter Flying Qualities, Requirements for
3.7 Vibration Characteristics
3.7.1 In general, throughout the design flight envelope, the helicoptershall be free of objectionable shake, vibration, or roughness. Specifi-cally, the following vibration requirements shall be met:
(a) Vibration accelerations at all controls in any direction shallnot exceed 0.4g for frequencies up to 32 cps and a doubleamplitude of 0.008 inch for frequencies above 32 cps; thisrequirement shall apply to all steady speeds within thehelicopter design flight envelope and in slow and rapidtransitions from one speed to another and during transitionfrom one steady acceleration to another.
(b) Vibration accelerations at the pilot, crew, passenger, andlitter stations at all steady speeds between 30 knots rearwardand VC . shall not exceed O.15g for frequencies up to 32cps an UlSaouble amplitude of 0.003 inch for frequencies greaterthan 32 cps. From Vc . to VLi . the maximum vibratory accel-eration shall not exceeh 6.2g up o 36 cps, and a double ampli-tude of 0.003 inch for frequencies greater than 36 cps. At allfrequencies above 50 cps a constant velocity vibration of 0.039fps shall not be exceeded.
(c) Vibration characteristics at the pilot, crew, passenger, andlitter stations shall not exceed 0.3g up to 14 cps and a doubleamplitude of 0.003 inch at frequencies greater than 44 cps dur-ing slow and rapid linear accelerations or deceleration fromany speed within the design flight envelope.
3.7.2 The magnitude of the vibratory force at the controls in any direc-tion during rapid longitudinal or lateral stick deflections shall not ex-ceed 2 pounds. Preferably, these vibratory forces shall be zero.
3.7.3 The helicopter shall be free from mechanical instability, includ-ing ground resonance, and from rotor weaving and flutter that influencehelicopter handling qualities, during all operating conditions, such aslanding, takeoff, and flight.
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MIL-STD-810B(4) Environmental Test Methods
4.5 Common test techniques.-
4.5.1 Sinusoidal vibration tests. - The vibration shall be applied
along each of the three mutually perpendicular axes of the test item.The vibratory acceleration levels or double amplitudes of the specifiedtest curve shall be maintained at the test item mounting points. Whenspecified, for sinusoidal resonance search, resonance dwell, and cyclingtests of items weighing more than 80 pounds mounted in airplanes, beli-copters, and missiles, the vibratory accelerations shall be reduced +/-1g for each 20 pound increment over 80 pounds. Acceleration derating shallapply only to the highest test level of the selected curve, but in nocase shall the derated test level be less than 50 percent of the selectedcurve (see note 1 of applicable table 514.1-1 through 514.1-V). Forequipment weighing over 100 pounds and transported by aircraft, resonancesearch, resonance dwell, and cycling tests may be frequency and accelerationderated (see notes 1 and 2 of table 514..-VII). When packaged items arealways grouped together on mechanized loading platforms or pallets, accel-eration and frequency derating may be based on the total load on the pallet.When the input vibration is measured at more than one control point, thecontrol signal shall be the average of all the accelerometers unless other-wise specified. For massive test items, fixtures and large force exciters,it is recommended that the input control level be an averag of at leastthree or more inputs.
4.5.1.1 Resonance search. - Resonant frequencies of the equipment shallbe determined by varying the frequency of applied vibration slowly throughthe specified range at reduced levels but with sufficient amplitude toexcite the item, Sinusoidal resonance search may be performed using thetest level and cycling time specified for sinusoidal cycling test, providedthe resonance search time is included in the required cycling testtime of4.5.1.3.4.5.1.2 Resonance dwell -The test item shall be vibrated along each axis
at the most severe resonant frequencies determined in 4.5.1,1. Test levels
frequency ranges, and test times shall be in accordance with the applicable
conditions from tables 514.1-1 through 514.1-V figures 514.1 through 514.1-7 for each equipment category. If morc than four significant resonant fre-
* quencies are found for any one axis, the four most severe resonant frequen-cies shall be chosen for the dwell test. If a change in the resonant fre-quency occurs during the test, its time of occurrence shall be recordedand immediately the frequency shall be adjusted to maintain the peak reson-ance condition. The final resonant frequency shall be recorded.
4.5.1.3 Cycling - The test item shall be vibrated along each axis inaccordance with the applicable test levels, frequency range, and timesfrom tables 514.1-I through 514.1-VII and figures 514.1-1 through 514.1-7.The frequency of applied vibration shall be swept over the specified
123
range logarithmically in accordance with figure 514.1-10. The specifiedsweep time is that of an ascending plus a descending sweep and is twicethe ascending sweep time shown on figure 514.1-10 for the specified range.Linear sweep rates may be substituted for the logarithmic sweep rate.When linear sweep rates are used, the total frequency range shall bedivided into logarithmic frequency bands having similar time intervalssuch that e ach time interval is the time of ascending plus a descendingsweep for the corresponding band. 'he sum of these time intervals shallequal the sweep time specified for the applicable frequency range. Thelinear sweep rate for each band is then determined by dividing each band-width in cps by One-half the sweep time in minutes for each band. Thelogarithmic frequency bands may be readily determined from figure 514.1-10.The frequency bands and linear sweep rates shown in table 514.1-IX shallbe used for the 2 (or 5) to 500 cps and 5 to 2,000 cps frequency ranges.For test frequency ranges of 100 cps or less, no correction of the linearsweep rate is required.
.2I
uI,
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