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307
VIBRATION EFFECTS ON HELICOPTER RELIABILITY 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-
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_______Jl___IIII _
AD
USAAMRDL TECHNICAL REPORT 73-11 0VIBRATION EFFECTS ON HELICOPTER
RELIABILITY AND MAINTAINABILITY By Angelo C.Veca
April 1973
EUSTIS DIRECTORATE U.S.ARMY AIR MOBILITY RESEARCH AND DEVELOPMENT LABORATORY FORT EUSTIS, VIRGINIA CONTRACT DAAJ02-71-C-0037 SIKORSKY AIRCRAFT DIVISION UNITED AIRCRAFT CORPORATION STRATFO!?D, CONNECTICUT
Approved for public release; distribution unlimited. Rop-d.ced by
NATIONAL TECHNICAL
INFORMATION SERVICE 5
044, VA 221It
DISCLAIMERS The findings in this report are not to be construed as an official Department of the Army position unless so designated by other authorized documents. When Government drawings, specifications, or other data are used for any purpose other than In connection with a definitely related Government procurement operation, the United States Government thereby incurs no responsibility nor any obligatic. i whatsoever; and the fact that the Government may have formulated, furnished, or in any way supplied the said drawings, specifications, or other data is not to be regarded by implication or otherwise as in any manner licensing the holder or any other person or corporation, or conveying any rights or permission, to manufacture, use, or sell any patented invention that may in any way be related thereto. Trade na m es cited in this report do not constitute an official the use of such commercial hardware or software.
endorsement or approval of
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IS*e.witv cleam D0hation,of lots, bodt, of .astrot Iend irdo ine arnotatio.. must be entems! whon the ... rall twpoef to r.saviflest ,ORIGINATING ACTIVITy (COWI~.tO eUtfti)U. REPORT SECURITY CLASSIFICATION
Sikorsky Aircraft Division
United Aircraft Corporation
21
Stratford, Conaecticut IREPORT
dnlssfe classifie
TITLE
IVIBPATICN EFFECTS 0:1 HELICOPTER RELIABILITY AND MAINTAINABILITY UESCRIPTIVX NOTES (I'yps of report end inclusive dates)
4.
Final Report S A4TP4O.S)(Ji8FatM~
middle Inlitial. le*t Rem.)
Angelo C. Veca 6, ALPORT DATE
Tb.
7a. TOTAL NO. OF PAGES
Apil 1973
/2
e.CON-MACY OR GRANT NO0.
.
orO
Reps
5
0e. OMIGINA 1 09S McpGR r NUMUIRI)
DMJO02-71--0037
USAAMKDL Technical Report 73-11
A6 PROJECT No.
1F32oLDB38
______________________
C.
9b. DINER RE[PORT NORS) (An1y61hor numInhors Owl may be eeel&e1 O~le report)
Sikorsky Aircraft Report SER-611567
IL 10. O!STMIIUV*IOH STATCNT'
Approved for public release; distribution unlimited. 12- SPONSORING MILITARY ACTIVITY
11. SUPPLEIIENT1ARY NOTES
Eustis Directorate, U. S. Army Air Research & Development Laboratory, Fort Eustis, Virgi*.ia
IMobility 12. A ""T ACT
In this study, aifferenzea in reliability and maintainability data were examined on two groups of USAF 1-3 helicopters with distinctly different vibration characteristics. One H1-3 helicopter group was equipped with the rotor-mounted vibration absorber, a device which rvduces helicopter vibration induced by the rotor, and a second aircraft
group 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 rate and direct mainti-nance for the H-3 helicopters with absorbers and with reduced vibration levels. The overali H-3 helicopter failure rate and corrective maintenance are by 48% ant' 38.5%, respectively. The average reduction in vibration level was
freduced
ifed-zyz, coats show a siicf'wiuL reldutuiun or approximately ' 5~.%. c 10% for the overal. aircraft. At the subsystem and component levels, the same re ductions are shown in almost every case with the exception of certain navigation and avionics components. There are at least 2,zmilitary vibration specifications and standards which specify vibration criteria for design or test of airborne equipment. No obvious conflicts were found in these slecifications, but they are lacking in requirenents which clearly describe realistic vib-ation exposure times for the entire helicopter air vehicle system and its componerzts.
Few D D 110111OI1473
REPLACER! 00 FORM 947S. I JAM 04. WHICH in ONOIES
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UNCLASSIFIEDCesfceo
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UNCLASSIFIED Secb'ity'Clasmjllcalon 14.
LINK a
LINK A
Kery WORDS
ROL.9
WT
ROLIL
WT
LINK C ROLIC
WT
Helicopter Vibration Reliability Maintainability
Life-Cycle Cost
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UNCLASSIFIED !
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LINK A
LINK 9
KEY WORDS ROLE
WT
MOLE1
WT
Helicopter Vibration Reliability Maintainability Life-Cycle Cost
Secuty Claslsicatuic
LINK C ROLE
WT
Project IF163204DB38
Contract DAAJ02-71-C-0037 USAAR1RDL 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
i Prepared by Sikorsky Aircraft Division United Aircraft Corporation Stratford, Connecticut
for
Eustis Directorate U. S. ARMY AIR MOBILITY RESEARCH AND DEVELOPMENT LABORATORY FORT EUSTIS, VIRGINIA
Approved for public relcase;
distribution unlimited.
SUMMJARY This study assesses the effect of helicopter vibration environ.ent on helicopter subsystem reliability, maintainability, and life-cycle costs, and the adequacy of design and acceptance test specifications applicable to helicopter vibration. In this study, differences in reliability and maintainab.lity data were examined on two groups of USAF H-3 helicopters with distinctly different vibration characteristics. One H-3 helicopter group was equipped with the rotor-mounted bifilar vibration absorber, a device which reduces helicopter vibration induced by the rotor, and P.second aircraft group 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 rate and direct maintenance fo' the H-3 helicopters with absorbers and with reduced vibration levels. The overall H-3 helicopter failure rate 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% for the overall aircraft. At the subsystem and component levels, the same reductions are shown in eianost every case with the exception of certain navigation and avionics components. There are at least 2:9 military vibration specifications and standards which specify vibration criteria for design or test of airborne equipment. No obvious conflicts were found in these specifications, but they are lacking in requirerients which clearly describe realistic vibration exposure times for the entire helicopter air vehicle system and its components. As shown by this study, reduction in vibration levels can significantly improve reliability and reduce maintenance and life-cycle costs. The results also suggest that the useful life of an aircraft can be extended beyond current limits simply by reducing vibration exposure.
iii
FOREWORD The work for this study was authorized by Contract DAAJ02-71-C-0037, Project IIh'6320D.B38, issued by the Eustis Directorate, U. S. Army Air Mobility ana Development Laboratory, Fort Eustis, Virginia under the technical cognizance of Major A. Gilewicz.
LResearch
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 and Calculations. Fr. Michael Starzyk, Standards, Specification Aasessment. Mr. Spencer Lauer and Mr. Thomas Chernesky, Technical Computing, AFM 66-1 Data Reduction.
i
p
I
Preceding page bvii
fTABLE
OF CONTENTS
Page SUMMARY
.................................
FOREWORD ........
.1
......
..............................
LIST OF ILLUSTRATIONS ........ LIST OF TABLES ........
.........................
....
LIST OF SYMBOLS .........
viii
...........................
..
INTRODUCTION .........
v
...........................
....
xiii
............................
METHOD AND RESULTS .........
...
CONCLUSIONS .....
... ..
RECOMMENDATIONS .......... LITERATURE CITED ....
25
86 86 98
.....
106
........................... ......................... .........................
......................
5 5 9
...........................
Reliability ......... Maintainability .... ..... Specification Assessment .....
1
........................
Aircraft Population and Data Separation ...... ............. Determination of Vibration Magnitudes .... ... .............. Reliability and Maintainability Data Source and Analysis ..... Individual Aircraft Reliability and Maintainability Comparison ......... . . . . . .......... ......... Life-Cycle Cost Model and Life-Cycle Cost Determination ..... .. Vibration Design and Acceptance Test Specification Assessment .......... ............................ .. DISCUSSION OF RESULTS ........
xi
106
.....
...
...
208 l1
.
12
...
...........................
113
...
..........................
l11
APPENDIXES
I. Climatic Briefs .... II.
......
..
.............
Excerpts From Assessed Specifications. . .....
DI$TRIBUTION
. . . . . . . . . . . .
Preceding page blank
vii
....
............ ...
. .
115 122
.
. 125
LIST OF ILLUSTRATIONS
Page
Figure 1
2
3
Fixed-Wing Electronic Reliability vs Rotary-Wing ...................... ..... Reliability .......
2
Fixed-Wing Hydraulic Reliability vs Rotary-Wing Hydraulic Reliability ....... .. ..................
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 Levels
16
...................
BL IO(LT) WL 181.5 .............. 8
........
Extrapolation, Interpolation Vibration g Levels
BL !0(RT) WL 181.5 .........
9
17
...................
Extrapolation, Interpolation Vibration g Levels
Sta. 250 and Sta. 300, WL 181.5 ...
.............
Extrapolation, Interpolation Vibration g Levels ..................... BL 0, WL 181.5 .......
10
11
12
13 14
....
18
....
19
Linear Extrapolation, Interpolation of Vibration g Level for Entire CH-3 Aircraft Without Vibration ....................... ..... Absorber .......
21
Linear Extrapolation, Interpolation of Vibration g Level for Entire CH-3 Aircraft With Vibration ........................ .... ..... Absorber
22
Vibration Magnitude Profile Without and With Vibration Absorber at Cabin Ceiling Level, WL 181.5
23
.
•
Vibration Magnitude at Cabin/Cockpit Floor Level,
24
................
WL 107....... 15
Location Grid for CH-3 Aircraft Components ...
16
Comparison of Total Average Failure Rate and MMH/KFH for Top 13 Aircraft Subsystems ...... .............
67
Comparison of Average Failure Rate and MMH/KFH for Selected Airframe Subsystem Components . ............
68
17
.......
61
viii
;"
'
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. I'
.[
|
'...
'll
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l
1
v
'
,
,.
-
..
. ,
-
i
2
....
-
Page
Figure 18
Comparison of Average Failure Rate and M4H/1KH for ........... ... 69 Selected Drive Subsystem Components ....
19
Comparison of Average Failure Rate and M.4H/TH for Selected Utility Subsystem Components ...........
70
20
Comparison of Average Failure Rate and M4H/KFH for ........ Selected Landing Gear Subsystem Components ..
71
21
Comparison of Average Failure Rate and I4H/KFH for
Selected Fuel Subsystem Components .... 22
..
............... 72
Comparison of Average Failure Rate and 1N4H/'H for
Selected Flight Control Subsystem Components ........ ... 73 23 24 25
Comparison of Average Failure Rate and N4R/KFH for Selected Cockpit/Fuselage Subsystem Components ...
...... 74
Comparison of Average Failure Rate and MMH/KFH for ............ Selected Electrical Subsystem Components ..
75
Comparison of Average Failurp Rate and MMH/KFH for ... Selected Hydraulic Power Subsystem Components .........
76
26
Comparison of Average Failure Rate and MMH/KFH for Selected Intercommunications Subsystem Components .......77
27
Comparison of Average Failure Rate and MMH/KFH for Selected Radio Navigation Subsystem Components ......
28 ,
29
30 31 32 33 34
78
Comparison of Average Failure Rate and MOIH/KFH for Selected Airconditioning/Heating Subsystem Components . . .
79
Comparison of Average Failure Rate and W.IH/KFH for ............. all Internal and External Lights ......
80
Comparison of Average Failure Rr.te and MMOH/KFH for ...................... all Switches ...........
81
Comparison of Average Failure Rate and MMH/KFH for ................ all Connectors/Plugs/Wiring .....
...
82
Comparison of Average Failure Rate and MMH/KFH for .................... all Hoses and Lines .......
...
83
Comparison of Average Failure Rate and NMH/KFH for ........................ all Valves ........... Comparison of Average Failure Rate and MMH/KFH for ...................... all Relays .......... ix
84
...
85
Page
Figure
35
Distribution of Individual Aircraft Failure Ra-tes Comparing Aircraft Groups Without and With Absorber . ..
87
Distribution of Individual Aircraft M.i/FH Comparing Aircraft Groups Without and With Absorber ... .....
88
37
Life-Cycle Cost Model ......
89
38
Vibration Specificazion Tree
36
................... ... ...............
x
100
LIST OF TABLES Table I II
III IV
V
VI
VII VIII
IX
X XI
XII XIII XIV XV XVI XVII XVIII
Page Selected H-3 Aircraft, Locations, and Times .... H-3 Aircraft Mission Profile ........
........ ..............
Aircraft Vibration Response .......
8 9
................
10
Total Aircraft System Comparison, Reliability and Corrective Maintenance ...... ..................
...
Component Comparison of Reliability and Corrective Maintenance at Subsystem Level ....... ............. Compcrnent Comparison of Reliability and Corrective Maintenance - Internal and External Lights ....
27
29
........
Component Comparison of Reliability and Corrective Maintenance - Switches ..... ..................
39
.. ..
Component Comparison of Reliability and Corrective Maintenance - Connectors/Plugs/Wiring .... ...........
45
Component Comparison of Reliability and Corrective Maintenance - Hoses/Lines ...... .................
47
Component Comparison of Reliability and Corrective Maintenance - Valves ...... ...................
.... 53
Componert Comparison of Reliability and Corrective Maintenance - Relays ...... ...................
.... 59
Individual Aircraft X and MH/KFH Without Absorber ........
63
Individual Aircraft A and MMH/KFH With Absorber ........
64
Comparison of Aircraft A and M24H/KFH by Population Group on Aircraft Action Only .... ..............
... 65
Comparison of Scheduled Maintenance Actions and ...................... 66 Maintenance Man-Hours ....... Life-Cycle Cost Model Assumptions ....
................ 91
Life-Cycle Savings per Aircraft Resulting From Vibration .Reduction........ ..................... Life-Cycle Model Nonbifilar Subsystem Spares and Abort Rates ......... ........................
94 ...
96
xi
". ,. "--,r %,L"... . "=:"': =' -
m" " -#'m= ",
- . . .. . ... . .. .. .... _ . 4
Page
Table XIX
XX
Life-Cycle Savings per Aircraft Resulting From ..... .......... . .. . Vibration Reduction ......
97
.
Applicable Specifications With Vibration Requirements
.
99
Ratio Change in Average Failure Rate and Ratio Change in Vibration Level ....... ...................... ..
107
Ratio Change in .2IH/FH and Ratio Change in Average Vibration Level ....... .........................
110
-XXIII
AWS Climatic Brief, Eglin AFB, Fla ...................
115
TXXIV
AWS Climatic Brief, Forbes AFB, Kansas .....
XXI
XYII
XXV XXVI XXVII XXVHII XXIX
AWS Climatic Brief, Parks AFB, P.I .....
..........
AWS Climatic Brief, Thule AFB, Greenland ..
11
.............
AWS Climatic Brief, Shaw AFB, S. Carolina .....
116
..
118
........... 119
AWS Climatic Brief, Wcodbridge Aerodrome, England .......120 Climatological Data for Elmendorf AFB, Alaska ..........
121
xi
xti
-- --
OF SYMBOLS
CLIST
A
lateral acceleration amplitude
B
vertical acceleration amplitude
DH/D
down hours per day
F
primary rotor excitation frequency, rad per sec
g
gravitational acceleration, 32.2 fT per sec
H
vibratory frequency, cycles per sec
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
n
2
z
Rtotal
vibration response
R
mean square ratio
RI
reliability sensitivity index
t
mission time, hr
T
test statistic
Vc
cruise speed, kt
Xlateral Zvertical
acceleration acceleration
AI
change in failure rate
A
failure rate:
A a
abort rate
failures per 1000 flight hours
xiii
INTRODUCTION Vibration has a recognized influence on the reliability and maintainability of helicopter airborne equipment. Airborne equipment failure rates - such as those associated with hydraulics, power train, structure, furnishings and flight controls - are expected to be related to the frequency, amplitude, and duration of the vibration environment. It is not readily apparent, however, whether or not this effect of vibration is highly significant or economically important. The methods available for predicting reliability and maintainability stress the importance of the environmental effects which may degrade the reliability of airborne equipment. To date, the prediction handbooks provide a constant multiplier factor K for ranges of environmental effects to 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 are affected to different degrees by each environment, more accuracy in the reliability prediction can be attained by developing failure rates around each environmental factor. Component failure rates in fixed-wing application are demonstrauly lower than those of equivalent or similar componentz used in helicopters. This observation is verified to some extent by comparing fixed-wing and helicopter failure rate- appearing in the Bureau of Nave Weapons Failure Rate Data Handbook.Figures 1 and 2 illustr-te the comparison for hydraulic actuators and selected electronic components. This trend suggests that lower vibration levels inherent in fixed-wing aircraft lead to better component reliability, but other major differences in the applications alo exist. At the same time, however, significant reductions in maintenance have been reported on commercial S-61 model Delicopters when they were equipped with vibration-reducing equipment. The deleterious effect of vibration on reliability is not readily separated from other environmental effects. Starting in 1970, however, a unique opportunity to relate helicopter vibration data to field reliability data became possible with the installation of rotor-mounted vibration absorbers on USAF Ct{-3 helicopters. Recorded reliability data and measured vibration data on these type aircraft before and after the installation of the absorber were available. The measured vibration data were acquired from a test program conducted on three CH-3 helicopters at Sikorsky Aircraft. Reliability data were acquired on operational aircraft currently in the field from the USAF AFM4 66-1 maintenance data reporting system. The study reported herein evaluated vibration levels and reliability and maintainability records for CH-3 helicopters with and without the vibration absorbers installed. Two CH-3 helicopter populations consisting of 15 aircraft each were selected for this study. One aircraft group was initially placed into service with the absorber and the other group was placed into service without the absorber; each aircraft group had
1
1500-
1000 in I
C-
z 0
F-
w _Jw
500-
FIXED WING
HEL I COP T ER 0
I
I
I
50
100
150
RELATIVE COMPL EXITY Figture 1-
Fixed-Wirg Electronic Reliability vs
iotriry-Winp, Reliability.
35-
30-
C 25FIXED WING
X
U20-
0 i15F0
HELICOPTER
10-
0
1
2
3
4
RELATIVE COMPLEXITY Figure 2. Fixed-Wing Hydraulic Reliability vs Rotary-Wing Hydraulic Reliability.
approximately the same number of flight hours for the time period covered by the study. The vibration data and reliability data were generated and recorded prior to commencement of this study. Changes in CH-3 R/M levels due 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 entered service equipped with the vibration absorber. Thus, a corresponding group of 15 H-3 helicopters were selected as a control group without the absorber. Figures 3 and 4 illustrate the helicopter and the vibration absorber respectively. The AFM 66-1 data received from 1.PAFB contain information on all H-3 aircraft in service, and the computer programs at Sikorsky Aircraft were modified to allow the extraction of those data which would be applicable to the study. For the "without-vibration-absorber aircraft," data were taken from 15 aircraft serial numbers selected from the last group of aircraft delivered without the absorber. Only data prior to January 1970, the date of initial delivery of vibration absorber kits to the USAF, were used. Information for the "with-vibration-absorber aircraft" was taken, by aircraft serial number, only from aircraft delivered with absorbers installed. The 2light-hour totals for each aircraft were determined from the aircraft logs or from flight times prcvided by field service reports. In conjunction with the determination of total flight time, the time spent in performing various missions was also considered since reported reliability and maintainability data may be seusitive to the particular mission. Detailed mission profiles were obtained for the two fleets, and no significant differences were found in the way, purpose, and length of time the aircraft were being used. Therefore, relisbility and maintainability sensitivity to missions flown vas assumed to be equivalent for both groups of aircraft. The total number of flight :aours accumulated by each aircraft over the 14-month period covered by the AFM 66-1 data along with aircraft locations are provided in Table I. The percentage utilization for each mission is presented in Table II. The geographical locations of the sample groups of aircraft suggest that they may have been exposed to large differences in climatic conditions and that this would impact uDon the reliability and maintainability data studied. The aircraft without the absorber are located in geographical regions ranging from the Tropic Zone to the Temperate Zone (l14ON Lat to 520N Lat). The aircraft with the absorber are located in geographical regions ranging from the northern portion of the Tem~erate Zone to a region above the
Arctic Circle, the North Frigid Zone (61 N Lat to 76011 Lat).
Climatolog-
ical surveys were investigated and the aircraft without the absorber were exposed to mean daily temperetures ranging from 40oF to 890 F, mean annual snowfall ranging from 10 inches to 23 inches, and mean wind speed ranging from 6 kt to 9 kt. The aircraft vith the absorber were exposed to mean daily temperatures ranging from 5OF to 42.6 0 F, mean annual rainfall of 5.5 inches to 17 inches, mean annual snowfall ranging from 10.5 inches to 36 inches, and mean wind speeds ranging from 5 kt to 7 kt.
5
r- w
-~
--
-
-----.
---.-------------
-
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~---~-~
A. C
w.
I-
'4
2.
/ 6
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-
i 4
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N N
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~1.
4 7
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I.
tTABLE
SELECTED H-3 AIRCRAFT,
LOCATIONS, AND TIMES
~Accumulated Aircraft Serial SNumber Without Absorber
Flight Hours (Utilization) (3/68
-
Date Entered Service
Location
4/69)
66-13284
525
6/67
Eglin AFB, Fla.
-13285 -13286
530 557
7/67 7/67
Forbes AFB, Kan. Eglin AFB, Fla.
67-14705 -14707 -14711 -14713
520 595 684 725
12/67 6/68 3/68 3/68
Forbes AFB, Kan. Shaw AFB, S.C. Eglin AFB, Fla. Eglin AFB, Fla.
-14714 -14715 -14716 -14717 -14719
559
4/68 4/68 5/68 5/68 7/68
Eglin AFB, Fla. Woodbridge, G.B. Woodbridge, G.B. Woodbridge, G.B. Forbes AFB, Kan.
-14720 -14723
476 18o
8/68
9/68
Shaw AFB, S.C. Clark AFB, P.I.
-14724
85
10/68
Clark AFB, P.I.
Total Hr
6228
With Absorber
116 125 126 425
(1/70
-
4/71)
69-5798
490
4/70
Thule AFB, Greenland
-5789 -5800 -5801 -5802 -5803
510 511 550 445 405
4/70 4/70 4/70 5/70 5/70
Thule AFB, Greenland Alaskan Air Com. Alaskan Air Com. Alaskan Air Com. Alaskan Air Com.
-5804 -5805
500 495
6/70 6/70
Alaskan Air Com. Alaskan Air Com.
-5806
450
7/70
Alaskan Air Com.
-5807 -5808 -5809 -5810 -5811 -5812
415 335 345 300 230 190
7/70 8/70 9/70 10/70 11/70 1/71
Total Hr
6171
8
Alasken Alaskan Alaskan Alaskan Alaskan Alaskan
Air Air Air Air Air Air
Com. Com. Coi. Com. Com. Com.
TABLE II.
H-3 AIRCRAFT MISSION PROFILE
Percentage Utilization
Mission
33.6
Training
9.6
Rescue
h8.8
Logistics
8.0
Other
The climatological summaries for each location cited in Table I are contained in Appendix I. The ranges of climatological conditions cited above are within the specified range of values for the H-3 aircraft. The data accumulation period for each population extended over a period of 14 months (more than a full year), so that it cannot be argued that the absorber aircraft benefited by postponing maintenance from winter months to the more favorable summer months. It appears that the climatic differences 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 and maintainability would be biased in favor of the aircraft without the vibration absorber and may reduce the apparent effect of lower vibration levels on reliability and maintainability values for aircraft with the absorber. The conclusion of this report as to the beneficial effect of vibration reduction on improving reliability and maintainability may therefore be conservative when considering the climatic difference.
l,
DETERMINATION OF VIBRATION MAGNITUDES The measured vibration data were acquired from a test program conducted on three H-3 helicopters at Sikorsky, whereas the reliability data were acquired on operational aircraft currently in the field from the USAF AFM 66-1 maintenance data reporting system. The question arises as to the rigor of using vibration data taken from aircraft different from those from which the reliability data are taken and pooling these data to form the basis of this study. It has long been an established procedure in the aircraft industry to acquire various data on a sample of aircraft and to apply the results of these data throughout the entire fleet of aircraft. Vibrations are induced in the helicopter and its components by the main rotor at a frequency = Hz or Fn/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 results recorded on H-3 helicopters both with and without a bifilar vibration absorber. These tests measured the vertical and lateral vibration amplitudes and directions at various points through the aircraft at the dominant frequency of 17 H z , and values obtained are shown in Table III (as extracted from the test report). TABLE III. AIRCRAFT VIBRATION RESPONSE* Without Absorber Butt Water Station Line Line (in.) (in.) (in.)
Vertical Lateral Accel Accel (g)
(g)
With Absorber R
Vertical Accel
(g67
()
Lateral Accel (g;
ax
(..
95 95
21(RT) 107 21(LT) 107
0.17 0.20**
0.31 0.31
0.35 0.31
0.11 0.09
0.29 0.29
0.31 0.30
187 187
39(RT) 107 39(LT) 107
0.17 0.17
0.22 0.22
0.28 0.28
0.09** 0.17
0.16 0 16
0.16 0.23
243 243 290 N90 379 379
39(RT) 39(LT) 39(RT) 39(LT) 39(RT) 39(LT)
0.22 0.24 0.19 0.13 0.25 0.17**
0.32 0.32 0.33 0.33 0.23 0.23
0.39 0.40 0.38 0.35 0.34 0.23
0.05 0.16 0.35 0.16 0.15 0.05**
0.19 0.19 0.13 0.13 0.17 0.17
0.20 0.25 0.14 0.21 0._3 0.17
243
10(RT) 181.5
0.39
1.16
1.22
0.23
0.42
0.49
243
10(LT) 181.5
0.77
1.16
1.39
0.35
0.42
0.55
290
10(RT) 181.5
0.75
1.34
1.54
0.25
0.45
0.51
290 542
l0(LT) 181.5 0 16o
0,56 0.24**
1.34 0.90
1.45 0.90
0.15 0.13
0.45 0.57
0.47 0.58
0.19
1.90
1.90
0.24
0.94
0.97
709.5
0
107 107 107 107 107 107
225
*Vibration Frequency of Five Per Rotor Revolution
**
90
or 270 °
Since the aircraft systems and corponents are exposed to both the vertical and lateral motions simultaneously and experience the resultant effect, the lateral and vertical vibration components were combined to provide a single resultant vibration value. Longitudinal vibration is not included because past flight surveys have shown it to be negligible. The procedure used to determine the vibration response magnitude is provided by the 10
equations below. Given
(2)
Y = A Sin (Fnt
(3)
+
z = B Sin F t n
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 vibration results in equation (5) above and is valid for any phase angle €. The phase angles used in cal-uiating the values shown in Table III were taken directly from the flight test data to the nearest 90 ° . Setting the time derivative of eouation (5) equal to zero and substituting 0, 900, 1800, or 2700 results in equations (6), (7) and (8): or 1JO°
R
= (A2 + B2 )
lId
=A for A> B
@
900 or 2700
CO
IRImax = B for B> A
@
900 or 270'
(8)
max
0
@
(6)
max
Id
(the maximum one-half peak to peak value of The absolute value of the vibration level) is Wependent upon the relative phase angle 0. The equations (6), (7), and (8) were used to determine the resultant vibration levels snown in Table III. The phase angles for most locations are 00 and 1800 where equation (6) applies, and the locations where phase angles are 900 or 270 are noted by an asterisk. In these cases, the larger 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 lateral vibration components are given in Table III and are mapped schematically in Figure and in Figure 6 for the H-3 helicopter without and with the vibration 5 absorber respectively. Sample Calculation Case I: From equation (5),
00 or 1800 2 2 (A + B)
R
0 = F (A2 +B n
max
at Therefore,
Fn t =
WRa
giving
=
ma
Given:
cosF t n
(9) cos F t n
(10)
/2, 3u/2
(A2 + B2 )
(6)
Vibration Level at Station 95, Butt Line 21(RT), Water Line
107 =
RL
0.31g,00
0.17.gO 0°
Z
(Table III, without absorber)
0.359
Case II: From Equation (5),
= 90 0 or 2700 R = {A2 (cos2 Fnt) + B2(Sin2 Fnt)} o
dR
t
ax
=0
(A2 (Cos2 Fn t) + B (Sin F t)) A2
(B
Therefore, Giving:
Fn t ax
(11)
)
Cos F t Sin F t
(12)
= 0, n/2, n, 3n/2 = A if A>B
(7)
= B if B>A
(8)
12
LT. 40
RT. 40
BUTT LINE 0 20 20
LT. 40
BUTT LINE 20 20 0
RT. 40
STA. 50 COPILOT 0.31
PILOT 0.35
100 150
D.28 O.40 : .35
.28C
200
.39c
250
1.390
01.22
.38C
300
IA5O
01.54
WL
i181.5
350 .34I
:*.23
400
500 .90
550 WL
160
600
Li
650 (I.
*
L =225
750
160 AND AB VE WATER LINE CEILING LEVEL
107 WATER LINE FLOOR LEVEL
Figure 5.
Total Vibration Response g Without Absorber. J.3
LT. 40
RT. 40
BUTT LINE 20
LT 4Q20
BUTT LINE 20 0 20
RT. 40
STA.
50 COPILOT 0.30
PILOT 0.31
100
150 3.23
.16 C
20
0.25
.20(
250
.550
D.21
.14C
300
.470
WL
0.49 0.51
181.5
350 .17
.28 C
400
450
500 550
().58 W L =160
600 650 700
.97
WL =225
~750\ 107
WATER LINE FLOOR LEVEL
Figure 6.
160 AND ABOVE WATER LINE CEILING LEVEL
Total Vibration Response g With Absorber.
Given:
Vibration Level at Station 95, Butt Line 21(LT1M
Water Line
107 0.31L0 °
11
Y
0.20gL90 0
=4o.31g
(Table III, without absorber )
(13)
(14)
The overall vibration characteristics throughout the H-3 aircraft were defined from the 16 pairs of measured vibration data points by making a linear point-to-point interpolation or extrapolation (as the case demanded). To illustrate the procedure, the interpolation and extrapolation of the ceiling data points (water line 181.5) were carried out as follows for the without-absorber case: At butt line 10 left the l.| values were assumed to vary linearly through the value of !.39g at station 243 and the value of 1.45g at station 290 to yield the interpolated value of l.hOg at station 250 and an extrapolated value 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 243 and 1.54 at station 290 were assumed to vary linearly to yield an interpolated value of 1.27g at station 250 and an extrapolated value of 1.60 at station 300. This is illustrated in Figure 8. Curves were then p ised through the station 250 data points to provide an estimate of the highest value near butt line zero and to fall off with constant slope either side of butt line zero. This shaping of the curve is suggested by the data points obtained, which seem to show maximum amplitudes at butt line zero near the rotor source of excitation. The lower values at BL-40 are also consistent with the lower measured levels on the floor below, (Figure 5). This procedure yielded butt line zero values 1.35g at station 250 and 1.6 0g at station 300. This process is illustrated in Figure 9. The peak value of 1.6 0g, station 300, butt line zero, w.-_r line 181.5 near the main rotor station, the center of vibration excitation, was taken as the maximum value at the point, and the drop-off from 1.60g to l. 35g in going from station 300 to station 250 was assumed to continue at constant rate proceeding toward the nose of the aircraft. Proceeding from station 300 toward the rear of the aircraft, a straight-line drop-off from the maximum value of 1.6 0g toiR . value of 0.90g measured at station 5h2 was assumed. In pr.ceeding Turther toward the rear, a straight line with increasing values ofrIAL was assumed until the measured value of 1.9g was reached at station 75V5. This procedure is illustrated in Figure 10, and represents a logical way of connecting the limited number of data points with a continuous line. Similar reasoning was used to establish the 1 Lax values along the butt line
15
1.8 E) INTERPOLATED OR EXTRAPOLATED VALUESSMEASURED VALUES
1.6
I
--
I
w
BUTT LINE iOL___
1.2
350
300
250
200
STATION
Figure 7. Extrapolation, Interpolation Vibration g Levels BL 1O(LT) WL 181.5.
16
IL
1.
INTERPOLATED OR
0 EXTRAPOLATED VALUES A MEASURED VALUES *
1.6, Il
1.e
1.4 BUTT LINE IOR
1.2 L 350
300
250
200
STATION
vi
Figure 8. Extrapolation, Interpolation Vibration g Levels, BL IO(RT) WL 181.5.
17
2.0-1~ STATION 300
1.6
1.21 STATION 250_____
.8 80L
-
~ 0
~
-
40L
BUTT
Figure 9.
-
-
4CR
8CR
LINE
Extrapolation, Interpolation Vibration g Levels Sta. 250 and Sta. 300, WL 181.5.
18
II
2.4 -
BUTTI LINE 10 -WITHOUT ABSORBER
-J w
-> w
--
.8
0
200
400
600
800
STATION
Figure 10.
Extrapolation, Interpolatior. Vibration g Levels, BL 0, WL 181.5.
19
corresponding to 20 and 40 inches left and right of butt line zero at the ceiling water line, for the floor level, and for the without and with vibration cases. Subsequent to the establishment of IRim values at the floor and ceiling water line levels, a straight-line inerpolation was used to establish values for the water lines between or beyond the floor or ceiling levels. Figures 11 and 12 show the completed process for thr .ithout-absorber and with-absorber cases, as described above, for the floor level and ceiling level vibration magnitudes. Vibration profiles, Figures 13 and 14, were developed from Figures 11 and 12 in order to portray the relative vibration amplitudes without absorber and with absorber at the ceiling and floor and provide a before/after picture at these water lines. These illustrations portray relative magnitude of the resultant vibration of the vertical and lateral vibration. The direction of vibration at each station is not indicated. Ranges of stations, butt lines, or water lines are specified as the location for some of the components considered in the study. The linear assumptions as to the vibration magnitudes acting on the component were applied and evaluated so that the lowest and highest vibration level is shown 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 = 1.24 g + 0.10g = 1.3hg The same procedure is carried out for the with absorber condition. Case II: Component:
Vibration Level at two points (without absorber) Anticollision Light
20
-FLT 40
BUTT LINE 20 20
RT 40
LT 4Q
STA
BUTT LINE 20 0 20
RT 40
50
(31
0.31
0.29 0.30 ().28
0.28
100
.56 0.64
.64 0.61
.56
.30 0.30 031 150 FLOOR
.74 0.80
.86 0.78
0.70
.33 035
.35
.28 0.28 M.28
022
0100 01.08 0.96
250
)1.10
01.26 01.35 01.14
01.15
200
WL =107 ).40 0.40
).40 039 039
0.35
0.36 0.36
0.37 0,3
300
$.27
0.29
033
350
01.19 01.38 01.20
0.21
0.25
400 450
0108 01.26 01. 9
.31
/4 ).28
031 ..
035 )34 4
OIAI
00.97 500
1.60 01.37 01.08
1.14 o .98 0-2
5 50
.84 "CEILING W L 181.5 0.98
.
.92 4) 9
.90
92 9
.90
600 650 700
.0 W L =225
750
Figure 11.
Linear Extrapolation, Interpolation of Vibration g Level for Entire CH-3 Aircraft Without Vibration Absorber.
21
LT.
40
20
BUTT
0
LINE
20
RT.
40
STA.
0
LT
BUTT LINE
4020
1)
R.T
22.D40
50 30 0.30
. 0
031
0.26
.25
0.24 024
0.27
.31
I00
15
0.36
150
0.39
0.40
.36 037
.39
0.40
.37
).42
).23 022 0.20 0.18
0.16
200
0.25 0.24 0.22 0.21 FLOOR L= 107 .21 0.19 .17 0.16
0.20
250
.42 0.44 0.47' 0.43 039 CEILING W L0 = 11.5 )A6 0.52 (L52 0.46 .41
,.14
300
4O 043
.20
350
0.45
).20 e.22 023
400
0.45
0.18
0.18
.17
0.18 ,J7 .18
0.19
M0
0.20
.22
.23
.4 za-Q2"0
450
49 0.47 (.51
.7
0
550
*
600
0.46
).52 0. )4.).
47 500
).40
()540o.48 5
6
..56
.58, 7/
6()0).71 650
.84
7009 WL =225 750
Figure 12.
Linear Extrapolation, Interpolation of Vibration g Level for Entire CH-3 Aircraft With Vibration Absorber. 22
700
650
600
550
50)0 450
400
350
300
250 200
150
100
150
100
WITRUIT BIFILAR ABSL41BEP
70 650
600
550
5M0
450
400
350
300250
200
WITH BIFILAP. ABVVRBER
Figure 13.
Vibration Magnitude Prof'ile Without and With Vibration Absorber at Cabin Ceiling Level, WL 181.5. 23
FWD
450
400
350
300
250
150
200
100
WITPUT BIFILAR ABSORBER
14150, 1100-
350
30 WD 25r,
200
150
100
WiTH BIFILAR ABSORBER
Figure 14.
Vibration Magnitude at Cabin/Cockpit Floor Level, WL 107.
24
Locations:
Stations 250 and 720, Butt Line 0 - 0, Water Lines 80 and 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 = 0.08 g
The vibr-Ltion magnitude of the light at Station 720, Butt Line 0, Water Line 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 were obtained from the U. S. Air Force Maintenance Management System3 and contained the failure and maintenance data for the two groups of aircraft which are discussed in this report. These data were prepared and recorded by the U. S. Air Porce within the normal routine of aircraft operation and are considered complete to the extent required by USAF directives provided by Reference 3. These data were collected and recorded prior to the start of the study and were not specifically collected by the USAF in support of this study, nor were these data edited in any manner. Information pertaining to the aircraft discussed in the study has been extracted from the bulk data and listed in Table I covering a l4-month period of operation and representing 6228 flight hours and 6111 flight hours for the nonabsorber and absorber equipped aircraft respectively. The reliability and maintainability information contained in the AFM 66-1 tapes consist of date, job number, aircraft tail number, quantity of failures, action taken, when discovered, parts removed, how malfunctioned, man-hours, work performed on or off the aircraft, and work unit code number. Wr, unit codes identify preventive maintenance tasks as well as components which required corrective maintenance. The data were sorted by work unit code, quantity of failures, and maintenance man-hours for each aircraft subsystem and component. Because of the large number of discrete work unit codes assigned to the H-3 (approximately 2000), covering 37 general subsystem codes, the effect of vibration on reliability and maintainability on those items reflecting more than 10 to 15 failures within each general subsystem code for the ten subsystem codes reflectin6 the highest number of failures or maintenance man-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 to a particular aircraft tail number. The average failure rate for a given wo~k unit code was copputed by taking the ratio between the total number of failures recorded and the total accumulated flight hours on each sample group of aircraft. Similarly, the average maintenance man-hour per 1000 flight hours, 14H/KF{I, for a given work unit code was computed by taking the ratio between the man-hours recorded and the total accumulated flight hours for each sample group of aircraft. The result of this procedure is shown in Table IV at the subsystem level and 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 vibration reduction at the subsystem level on reliability and maintainability. The next procedure was to separate out from all systems, those classes of components 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 having failures recorded against them, regardless of the quantity of failures, were tabulated along with the computation of the average failure rate X and the maintenance man-hours per flight hour, MMH/KFH. The intent of this procedure was to determine if a behavior pattern of vibration with respect to reliability and maintainability could be recognized other than the dramatic differences in failure rate and maintenance man-hours per flight hours evidenced in Tables IV and V. These results presented in Tables IV, VII, VIII, IX, X, XI had no readily discernible characteristic other than that which is evidenced at the subsystem level. The locations or range of location for all components can be related to the actual aircraft by referring to the locating grid in Figure 15. The term "average failure rate" was generated because it was not possible to establish the type of failure distribution which fit the reliability data. A failure distribution could not be established because time-tofailure information was not included in the AFM 66-1 data. Past studies on the reliability characteristics on major components of the H-3 helicopter indicated an exponential distribution modified by early and wearout failure phenomena. However, since a constant failure or an early failure or wearout phenomenon could not be established for the data used, the term "average failure rate" is used.
26
TABLE IV.
TOTAL AIRCRAFT SYSTEM COMPARISON RELIABILITY AND CORRECTIVE MAINTENANCE -MMH/KH
Failure Rates (10-)I
i!
W/Out
With
Absorber
Absorber
MMH/KFH
115.9
592.3
209.7
382.6
61.1 50.3
371.8 I06.4 289.6 2h0.7 118.8 209.5 321.4
216.5 26.3 189.8
79.4 76.3 71.2
50.8 60.5 278.8 23.2 26.2 19.9 49.7
155.3 80.1 99.8 195.1 68.0 149.0 42.6 25.7 53.2 56.4 21.5
15.3
209.0
217.7
-8.7
8.8
95.7
36.1
59.6
16.6
11.8
94.2
88.6
2.4 36.2 6.7 17.6 2.9 4.7 0.2 0.2 9.14 40.4
10.3 8.3 8.2 5.5 5.3 4.0 1.7 0 -0.2 -0.4
15.9 125.9 69.3 67.9 21.9 13.4 4.3 0.2 38.8 163.7
1.4 107.4 33.5 93.1 12.3 9.3 0.3 0.3 36.4 188.2
Aircraft
W/Out
With
Subsystem
Absorber
Absorber
Airframe
223.7
107.8
Drive Utilities Landing Gear Lights Fuel Fit. Control Rotor Coukpit/Fus. Electrical Hyd. Power Inter Comm.
108.7 64.1 91.5 119.6 56.2 58.4 8o.4 33.1 35.6 37.1 39.5
47.6 13.8 44.8 29.3 22.8 22.8 51.0 9.9 12.4 17.1 21.2
Radio Nay.
65.5
50.2
Air Cond/Heat
27.1
18.3
Auto Pilot
28.4
Emer. Equip Aux Power Unit HF Comm. UHF Comm. IFF Misc. Comm. Weap. Del. Emer. Comm. VHF Radar Nav.
12.7 44.5 14.9 23.1 8.2 8.7 1.9 0.2 9.2 40.0
Failure
I
Rate
46.7 90.3 33.4 35.6 29.4 23.2 23.2 20.0 18.3
* Minus sign indicates an increase in rate.
27
48.9
45.6
5.6
14.5 18.5 35.8 -25.2 9.6 4.1 4.0 -0.1 2.4 -2h.5
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VIBRATION DESIGN AND ACCEPTAICE TEST SPECIFICATION ASSESSMENT The objective of this study task was to assess applicable vibration design and acceptance test specifications in light of the reliability and vibration data presented. The procedure used to perform the assessment consisted of examining the vibration reqairements cited in each of the specifications applicable to components and systems used on the H-3 aircraft. Twenty-nine documents were reviewed and are listed in Table XX.. These documents are generally applicable to all helicopters and are not necessarily unique to H-3 applicaticns. Three documents were singled out for detailed assessment (MIL.-H-8501(A), MIL-STD-810B, and MIL-STD-781B) because they specifically cite the vibration requirements for helicopters, methods for environmental tests such as vibration, and tests for reliability where a vibration environment is simulated. Excerpts from MIL-H-8501(A) and MIL-STD-801B are contained in Appendix Ii. (The information on reliability testing in MIL-STD-781B is quite extensive; reference a complete copy to support the ensuing discussion). The specification tree provided in Figure 38 illustrates the interdependency existing between various component vibration specifications as well as the paragraphs within the specification where the requirement is cited. Current specifications for the acceptance testing of a helicopter component's ability to endure vibration are inadequate. Components are tested according to fixed-wing oriented specifications by employing relatively high frequency vibration (above 25 Hz). By contrast, helicopters are high-amplitude, low-frequency machines, the predominant frequency of excitation being between 10 Hz and 20 Hz. Further, helicopters are developed according to specifications which address the maximum allowable vibration levels in the cockpit and personnel cabin areas only. Lastly, in some cases, inventory helicopters do not conform to the military specification concerning vibration. A model specification is written declaring the higher vibration levels acceptable. Consequently, it is understandable why helicopter components, most of which are installed outside the cockpit and personnel cabin areas, often suffer premature failures. Further, understanding of the influence of excitation frequency on failure rate is needed before more specific recommendation can be made regarding military specification changes.
98
TABLE XX.
APPLICABLE SPECIFICATIONS WITH VIBRATIO,
:EQUIRE?_NTQ Title
Suecificatior M.LIL-H-8501(A)l IfL-S-8698(ASG)(l) MIL-T-8679 ?.IIL-D-23222A(AS) MIL-STD-8loB(4)
MIL-W-5013H(l) MIL-B-8584C
Helicopter flying & Ground Handling Qualities General Requirements For Structural Design Requirements, Helicopters Test Requirements, Ground, Helicopter remonstration Requirements for Helicopters Environmental Test Methods
Wheel and Brake Assemblies, Aircraft Brake System Wheel Design
MIL-T-50o1 F(l)
Tires, Pneumatic, Aircraft
MIL-L-8552C(2)
Landing Gear Shock Absorbers
MIL-F-18372(Aer)
Flight Control Systems Design Installation & Test Control & Stabilization Systems, Auto, Piloted Aircraft
MIL-C-18244A(Wep)
z.IL-T-6396C'ASG)
Tank, Fuel, Oil
MIL-T-5578C(2) MIL-F-17874B MIL-I-18802A(Wep)
Tank, Fuel, Aircraft Fuel System, Aircraft Fuel & Oil Lines, Aircraft, Instellation of Transmission Systems, VTOL-STJL, General Instruments & Nlavigation Equipment, Installation of Hydraulic Systems, Aircraft, Type I & II, Design, Installation Hydraulic System Components, Aircraft & Guided Missiles, General Cylinders, Aeronautical, Hydraulic Actuating, General Electrical Equipment, Aircraft, Selection & Installation Electrical Systems, Aircraft, Design & Installation Lights, Aircraft, General Specification Test Methods for Electronics and Electrical Components Electronic Equipment, Airborne, General Spec. Installation & Test of Electronics, Aircraft, General
I-IL-T-5955C MIL-I-18373A(AS)
1IL-.H-544OE MIL-H--8775C MIL-C-5503C(3) MIL-E-708OB(3) MIL-E-25499C
M.IL-L-6723B MIL-STD-202D(1) MIL-E-540OM(2) MIL-I-8700A
MIL-I-8677(Aer)
Installation Armament Control Systems
MIL-H-18325B(Aer) 14IL-STD-781B
Heating & Ventilation System, Reliability Tests: Exponential Distribution
99
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= = = =
reliability constant failure rate time varying failure rate time
In the reliability test methods of MIL-STD-781B, an attempt is made to not only subject the equipment to vibratory stress at a specified level, 2.2G ±10% peak, but also to include exposure time related to multiples of the specified MTBF. However, close examination of the test requirement shows that an equipment is exposed to the vibratory stress for approximately 1/6 of the total test tirae (hot/cold cycling occurring at the same time). Thus, MIL-STD-781B test methods also appear to be insufficient in simulating equivalent exposure time Lo vibratory stress since and equipment when airborne in a helicopter will be subjected to vibratory stress 100% of the time. Comparisons made between the vibration values cited in MIL-H-8501 (see Appendix II) and MIL-STD-810B or MIL-STD-781B show a difference in value which varies by an order of magnitude; i.e., 0.2g to 0.4g versus 2.Og to 5.0g. It is apparent that general design and test procedures related to vibration are not necessarily designed with reliability in mind. The discrepancy noted in MIL-STD-.781B, relative to exposure time, shows up a weakness in the reliability test specification which requires corrective action. 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 exposure times (other environmental influence notwithstanding). The above observations suggest that components specifications shoull relate more closely to the vibration environment that prevails in helicopters. The above observations also suggest that the allowable vibration specification for the helicopter (MIL-H-8501) should deal with the whole aircraft and not just the crew and passenger compartments. By having a more favorable vibration environment throughout the aircraft, component R/M would be improved. It is suggested that a happy meeting ground exists between making components more tolerant of vibration and making the whole helicopter less of 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 of vibratory stress. In general, a component subjected to a very large number of vibratory stress cycles will require a small percentage improvement in strength or reduction in vibratory stress to to from a severely limited life to an acceptable period of useful life. Many components exhibited a large percentage reduction or zero failures in the absorber-equipped aircraft as compared to a significant number of f.ilures in the nonabsorber aircraft. Table XXI compares the percentage chiange in the reliability to the percentage change in the average vibration level for the 13 subsystems considered in the study. The manner in which the one ratio is related to the other is unknown. The overall data presented in Tables V through XI, and Figures 16 through 34 suggest the response of reliability to changes in vibratory stress will possess different slopes for different types of components, dependent upon their construction, material used, method of installation, and location in the aircraft. Overall changes in the entire aircraft reliability with respect to changes in vibration can be observed from Table IV. There was a 48% reduction in the total aircraft failure rate between the H-3 helicopter without and with the vibration absorber. The evidence indicates, in all comparisons made, i.e., component level, subsystem level, and aircraft level, a decreasing failure rate with decreasing vibratory stress level. Several componerits listeA in Tables VI through XI did, however, show a somewhat higher failure rate for aircraft with the absorber. This effect is prevalent primarily in components located in the forward portion of the aircraft. Prior to adding the bifilar absorber to the 11-3, a battery absorber, momted in the forward equipment bay, was used to dampen vibrations in the cockpit area. (The battery absorber is standard equipment in the without-absorber group of aircraft). The vibration amplitude in the nose o" 'he aircraft was nearly the same for both aircraft populations. The battery absorber is very effective in reducing the vibration in the nose of the aircraft but is ineffective throughout the rest of the aircraft. Because the bifilar absorber had a greater overall effect on reducing aircraft vibration, the battery absorber was removed from aircraft having the bifilar absorber. Vibration data shown in Figure 11 are taken from aircraft equipped with the battery absorber, and the vibration data shown in Figure 12 are taken from aircraft with the bifilar absorber installed and battery absorber removed. (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.
System
RATIO CHANGE IN AVERAGE FAILURE RATE AND RATIO CHANGE IN AVERAGE VIBRATION LEVEL 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
107
-O,40
acquire proper tuning and damping factor). The heater ignition unit of the absorber equipped aircraft also showed an inordinate increase in failure rate, and it is suggested that the increase was caused by the requirement to use the heater more frequently in the northern latitudes where the absorber-equipped aircraft are located. The basic objective of showing impact of vibratory stress on the reliability characteristic of airborne equipment has been met and the tabulated and illustraed results strongly suggest that the reliability improves significantly when significant reductions in vibratory stress are achieved. The data suggest that the useful life of an aircraft can be extended without the need to strengthen or redesign certain airborne components to withstand vibratory stress if adequate methods of damping vibration or isolating equipment from v" atory stress are used. This assumes, however, that adequate testing .,ich simulates the actual vibration environment is also conducted. MAINTAINABILITY Corrective maintenance performed on an aircraft is a direct function of the reliability inherent in the design, and it follows that for any improvement made in relia ility a proportionate reduction in maintenance should also be achieved. The data analyzed in this study show that in all but a few cases drastic reductions in z.intenance were evidenced as a direct result of the reduced vibratory stress and the increased reliability resulting therefrom. However, the reduction in average failure rate does not fully account for the reduced maintenance because in addition to a lower frequency there is also a lower mean-time-to-repair in some cases. For instance, using data from Table V the average maintenance man-hors jer failure expended against the central frames assuming the same size repair crew, is given by M11H Failure
MH/!OOOFH Failures/lO00FH
(20)
For the without absorber case then MMH Failure
83.2 5.0
-.
6.0
(21)
and for the with absorber case then -
Failure
6.1 = 2.3
(22)
2.6
It appears that the extensiveness of the damage incurred (chargeable as a failure) is less in the with-absorber case than in the without-absorber case, and thus less repair work is required. This also implies, that although airborne components continue to fail or require corrective action,
108
the degree to which a component function is degraded and the extent to which repair time is required to return the component to a functional status are also reduced due to the improved vibration environment. This observation can be borne out by applying equation (20) to the data presented inTables VI through XI. Table XXII ;resents the ratio of change in 1,P.'/FH and the ratio of change in average vibration. The manner in which the one ratio is related to the other is unknown. The overall data presented in Table V through XI and Figures 16 through 34 suggest the response of mai.atenance to changas in vibratory stress will possess different slopes for lifferent types of components, dependent upon th2ir configuration and location in the aircraft. The study also considered the impact of the reductiou in vibration level on the preventive maintenance tasks. There was an increase in frequency and maintenance man-hours for preflight inspection and an increase in maintenance man-hours for postfl-ght inspections in the absorber-equipped aircraft. The reasons why this should occur in the face of reductions shown for corrective maintenance are unknown, but either local preventive maintenance policies or a more rigorous operational schedule may account for this effect. Table XV shows that although there was an increase in frequency and maintenance man-hours, the average number of maintenance man-hours per preflight inspection are 5.3 ,24H in the without-absorber case and 5.0 N24H in the withabsorber case. This implies a greatar aircraft operational frequency because prefligh+ inspections are done prior to the start of each flight. The 50% increase in the maintenance man-hours for postflight inspection in the absorber case is felt to be caused by local maintenance policy which allows portions of the periodic inspection to be performed when a postflight inspection is performed.5 This acccunts, in part, for the significant change in periodic inspection frequency and maintenance in the absorber-equipped aircraft. However, it is also known that the USAF changed the periodic inspection interval from 50 to 100 hours during the time period coverel by the data, sc a large portion of the reduction in frequency of 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 discovcrcd failures; thus, preventive maintenance intervals should be guided by the amount of fixing required subsequent to an inspection. If, in the long term, inspections do not turn up discrepancies, preventive maintenance resources should be conserved by increasing the inspection interval.
109
TABLE
XXII.
RATIO CHANGE Ili -Mi/FH AN!D RATIO CHA11GE IN AVERAGE VIBRATION LEVEL
SYSTEM
... F w i-!.r.1FHwl °
g
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
Airconditioning/Heating
.32
-.
.63
110
.40
SPECIFICATION ASSESSM
IT
The specifications governing the requirements for designing to a stated vibration environment are suitable in terms of determining resonant responses in an equipment, early fatigue failures, or gross deficiencies in mechanical design, but are lacking in requirements related directly to reliability. This is -,,ident even in the one specification written specifically for reliabili-ty testing, MIL-STD-781B. A more deliberate and well-designed specification along with detailed procedure is necessary such tha'u (1) helicopter vibration levels are speciied throughout the aircraft, (2) component specifications are related to the appropriate specified helicopter vibration levels, (3) the nature of the vibration level/endurance characteristics of different types of components are learned, a&d (M0 the statistical variability to be expacted among parts is taken into accouzit. The change in reliability and maintainability under the influence of vibration can be dramatic, as is evidenced by the data presented earlier, .nd thus, it becomes important that design and testing for the vibration environment be carefully scrutinized prior to planniug and performing reliability or maintainability demonstration tests. Considerable work has been performed and documented relative to developing vibration stress cycling curves for the many materials used in airborne aplications. However, except for special programs, little has been accomplished in the reliability area for complete equipment by type, function, or degree of complexity. For example, the difference in the average failure rates as related to difference in vibration level for the types of components shown in Tables VI through XI could not be generalized in a series of curves of mathematical expression because of the lack of data between the recorded points. However, this gap in the data could be filled by subjecting a large group of specimens to a test program which varies the vibratory strers over the range not covered. This kind of program would allow 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 a reduction in vibratory stress, indicates improvements of 48.7% in reliability 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 that there are variations in the way in which reliability and maintainability change with respect to change in vibratory stress even within families of similar components. It is concluded that more discrete data, acquired from a closely controlled test program, would be required on each group of components to determine the precise characteristic of the variation of reliability with respect to changes in vibratory stress. The reduction in the frequency of failure with respect to the improvement in the vibration characteristic suggests that the useful life of aircraft components can be extended without making any design changes in the equipment 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 in a component's life characteristic being changed from a severely limited life to a reasonable life. The improved reliability resulting from the reduced vibratory stress environment results in less corrective maintenance being expended on the CH-3 aircraft. This results in less downtime on the aircraft, thereby improving availability and contributing to the reduction in the operating cost of the aircraft.
The life-cycle cost analysis, based upon the data presented, shows that LCC may be reduced by as much as 10% for the 13 aircraft subsystems considered in the study, because of the improved reliability and maintainability brought about by diminishing vibratory stress. The reductions are manifested by lessening the costs of direct maintenance manpower and spares, and by improving helicopter utilization. Assessment of the various vibration design and test specifications reveal that they are inadequate relative to vibration design requirements and test criteria that can be related to the predictiun of actual operational reliability. The inadequacies result from insufficient knowledge of the vibration level/endurance characteristics of various types of aircraft components and from the lack of vibration level requirements for helicopter zones other than the cockpit and personnel cabin areas.
112
RECOIENDATIONS The following recommendations are made based upon the results of this study: 1.
Establish a vibration test program such that basic data can be acquired which will eventually allow the formulation cf the general relationships between reliability and effect of vibicatory stress 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 test program to expand the test to more complex or differently configured components.
2.
Ii3.
Perform basic research and analyses in order to establish an adequate specification which uniquely relates helicopter component reliability to vibratory stress exposure. Establish astudy/test pormto research and rptuonthe various devices, both active and passive, which will reduce the vibratory environment of helicopter components.
4. Expand helicopter vibration specifications to cover all significant areas of the aircraft (not just those occupied by personnel) where component fatigue damage may result.
!I
11
.1
113
LITERATURE CITED
-
I
1.
Naval Fleet Missile Systems and Evaluation Group, Failure Rate Data Handbook, Corona, California, September 1, 1969, Vol. lB.
2.
Paul, William F., DEVELOPM.ENT AND EVALUATION OF THE MAIN ROTOR BIFILAR ABSORBER, Sikorsky Aircraft Division of United Aircraft Corporation, 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 Force Base, Georgia.
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APPENDIX II EXCERPTS FROM ASSESSED SPECIFICATIONS MIL-H-8501A(l) 3.7
Helicopter Flying Qualities, Requirements for
Vibration Characteristics
3.7.1 In general, throughout the design flight envelope, the helicopter shall be free of objectionable shake, vibration, or roughness. Specifically, the following vibration requirements shall be met: (a) Vibration accelerations at all controls in any direction shall not exceed 0.4g for frequencies up to 32 cps and a double amplitude of 0.008 inch for frequencies above 32 cps; this requirement shall apply to all steady speeds within the helicopter design flight envelope and in slow and rapid transitions from one speed to another and during transition from one steady acceleration to another. (b) Vibration accelerations at the pilot, crew, passenger, and litter stations at all steady speeds between 30 knots rearward and VC . shall not exceed O.15g for frequencies up to 32 cps an UlSaouble amplitude of 0.003 inch for frequencies greater to VLi. the maximum vibratory accelthan 32 cps. From Vc . eration shall not exceeh 6.2g up o 36 cps, and a double amplitude of 0.003 inch for frequencies greater than 36 cps. At all frequencies above 50 cps a constant velocity vibration of 0.039 fps shall not be exceeded. (c) Vibration characteristics at the pilot, crew, passenger, and litter stations shall not exceed 0.3g up to 14 cps and a double amplitude of 0.003 inch at frequencies greater than 44 cps during slow and rapid linear accelerations or deceleration from any speed within the design flight envelope. 3.7.2 The magnitude of the vibratory force at the controls in any direction during rapid longitudinal or lateral stick deflections shall not exceed 2 pounds. Preferably, these vibratory forces shall be zero. 3.7.3 The helicopter shall be free from mechanical instability, including ground resonance, and from rotor weaving and flutter that influence helicopter handling qualities, during all operating conditions, such as landing, takeoff, and flight.
122
MIL-STD-810B(4) 4.5
Environmental Test Methods
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 specified test curve shall be maintained at the test item mounting points. When specified, for sinusoidal resonance search, resonance dwell, and cycling tests of items weighing more than 80 pounds mounted in airplanes, belicopters, and missiles, the vibratory accelerations shall be reduced +/-1 g for each 20 pound increment over 80 pounds. Acceleration derating shall apply only to the highest test level of the selected curve, but in no case shall the derated test level be less than 50 percent of the selected curve (see note 1 of applicable table 514.1-1 through 514.1-V). For equipment weighing over 100 pounds and transported by aircraft, resonance search, resonance dwell, and cycling tests may be frequency and acceleration derated (see notes 1 and 2 of table 514..-VII). When packaged items are always grouped together on mechanized loading platforms or pallets, acceleration 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, the control signal shall be the average of all the accelerometers unless otherwise specified. For massive test items, fixtures and large force exciters, it is recommended that the input control level be an averag of at least three or more inputs. 4.5.1.1 Resonance search. - Resonant frequencies of the equipment shall be determined by varying the frequency of applied vibration slowly through the specified range at reduced levels but with sufficient amplitude to excite the item, Sinusoidal resonance search may be performed using the test level and cycling time specified for sinusoidal cycling test, provided the resonance search time is included in the required cycling testtime of
4.5.1.3. 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 514.17 for each equipment category. If morc than four significant resonant frequencies are found for any one axis, the four most severe resonant frequencies shall be chosen for the dwell test. If a change in the resonant frequency occurs during the test, its time of occurrence shall be recorded and immediately the frequency shall be adjusted to maintain the peak resonance condition. The final resonant frequency shall be recorded. 4.5.1.3 Cycling - The test item shall be vibrated along each axis in accordance with the applicable test levels, frequency range, and times from 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 specified sweep time is that of an ascending plus a descending sweep and is twice the 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 be divided into logarithmic frequency bands having similar time intervals such that each time interval is the time of ascending plus a descending sweep for the corresponding band. 'he sum of these time intervals shall equal the sweep time specified for the applicable frequency range. The linear sweep rate for each band is then determined by dividing each bandwidth in cps by One-half the sweep time in minutes for each band. The logarithmic frequency bands may be readily determined from figure 514.1-10. The frequency bands and linear sweep rates shown in table 514.1-IX shall be 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 linear sweep rate is required.
.2
uI,
124
I
