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UJNCLASSI FIED4331586AD -__
DEFENSE DOCUMENTATION CENTERFOR
SCIENTIFIC AND TECHNICAL INFORMATION
CAMERON STATION. ALEXANDRIA. VIRGINIA
UNCLASSIFIED
NOTIC: When governme or other dravns, opeaiofications or other data are ued for my purposother than in connection vith a definitely related
zerzmnt proc IeMe o"ertion, the U. 5.Oovermnt thereby incurs no zesponsibility nor anyobllption vhteoweri nd the fact that the Ooves-mt may have fommlated fltohrmt6 or in ow %wsupplied the said drw n&i, speciftoations, or otherdata Is not to be reprded by IMU atiou or other-viae in szr m r licensing the bolder or anyother person or crponrtion, or conveying my rltsor pezuiseton to manufttacrt, use or sell mypatented invention that my in my vey be relatedtbereto.
RTI).TI).83--1230 433586lw
STATIC TEST PROGRAM FOR CH47A HELICOPTER
ROPIRT I , CIINEIVFER
I..
TECHINICAI IXX't'MEINTANY HFPORT %o, B'ID.TlIR41.4MI0
I)ECE.M Hl.Blt 1961:
All I'iCE FlII(IIl' IYNAMICS LABORATORYRESEAR(;ll AND TIECIIN)I.-Y DIVISIO,
AIR FORCE SYSTEMS COMMANDWHI;HIT.PAMrEBS(N AIR FORCE RAS.. OIIIO
I'"4 N l4 lI4 T., %, %t 111% ft,,l'. Snl rill.
BestAvailable
Copy
NOTICES
When Government drawings, specifications, or other data are used forany purpose other than in onnection with a definitely related Oovernment'procurement operation, the United States Oovernment thereby incurs r.0responsibUity nor aw obligation whatsoever; and the fee that the Oovern-ment may have formulated turnished, or in am way supplied the soid draw-ings. specifications, or other data, Is not to be regarded by implcation orotherwise as in any manner loensling the holder or any 3ther person orcorporation, or conveying say rights or permAssion to manufacture, use,or sell any patented invention that may in any way be related thereto.
Qualified roqucsters may obtain copies of this report from the DefenseDocumentation Center (DC), (formerly ASTIA), Cameron Station, Bldg. 5,5010 W'ke Street, Alexandria, Virginia, 22314.
This report has been released to the Oftice of Technical Services. U.S.Department of Commerce, Washington 25. D. C.. in stock qantitieo forsale to the general public.
Copies of this report should not be returred to the Research aid Tech-nology Division, Wright-Patterson Air Force Base, Ohio. unless returnts requir-.d by security considoratlons. contrectual obligations. oron n specific document.
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RTrD-T)R -63-4230FORE WORD
This internal report wab prepared by the Air Force Piigiut Dynamics Laboratory an aformal record of rho structural test program conducted on the Vertol CH-47A helicopter.The structural tests reported were conducted by tbi Structures Division of the A ir ForceFlight Dynamics Laboratory at Wright-Patterso Air Aomc Bae, Ohio, with Mr. RobertL. Schneider acting as project: test engineer. Mr. Thoma P. Hughtes. senior project testengineer of the Structures Division aind Mr. Stanley SkIbo of the Vortol Division of RoeingAircraft Company assisted In the program. Moers. earl J. Hartzell ank f"halmar 1).Cruze were responsible for all Instrumentation.
This is the final report for the project.
RTD-TDR-63-4230
ABSTRACT
Results of structural tostn conducted on the complte ailrframe of the Vertol f1 -47Ahelicopter covering all critical flight, landing, lke-off, and ground handling conditionsare presented In this report. Two critical test conditions for growth potential are alsopresented.
The structure of the Verol CH-47A helicopter, with the exception of the litter and troopeot Installations, was capable of withstanding the static ultimate loads pnerated by 71different test conditions.
This technical documentary report has been reviewed and approved.
W. A. SLOAN. Jj .Colonel, USAFChief, Structures DivisionAP Flight Dynamics Laboratory
.. ; II'
RTD-TDR -63-42Y) 1-422
TABLE OF CONMTNlJV.
Pare
Introduction ...................................... 1 £jest Article and Loed Application Methods ................. I
Iln trumwfltaIti l ................................... 2 2
Decr lption of Structure .............................. 2 2
Test Conditlons, ate of Toots. andTsmRUtl ... ........ 3 3
Conclualon and Recommendations ...................... 9 9
APPENDIX - C-47A Structural Test Conditio ............. 1 I
iv
RTD-TDR -63-4230LIST OF ILLUSTRATIONS
Figure Page
Frontispiece Three Views of CH-47A Helicopter
I Structure Dead Weght Relief System ................. Is
2 Basic Test Set-Up ............................... 19
3 Basic Test Set-Up .............................. 20
4 Load Patches ................................. 21
5 Typical Air Bladders Used for Loading ................ 22
6 Typical Structure Used for Containing the LoadingBladders ............................... . ... . 23
7 Pallts Used fnr Floor l.oadlng ......... ........... 24
8 Typical Loading Straps Bonded by RTV SiliconeAdhesives ......................... .. ....... 25
9 CH-47A Structural Schematic .... ................ 26
10 Exploded View of CH-47A 1elicopter ................. 27
11 Failure of 114C1013-8 Link Assembly at85 percent [X . . .. ........ ...... .................... 28
12 C1-47A 114CI013-8 Link Assembly--Fallure at, 94 percen ..... 2......................... 9
13 Condition VB Aft Obstruction--Shear Failure InOutboard Shear Panel of Right Aft Gear Shear Bo ........ 30
14 Condition VB Failure In Outomard Sher Panel ofRight Aft Clear Shear Box ......................... 31
I5 Buckling of Part No. 114S3351 at Frame 440 BL-20 ....... 32
I fHuckling of Part No. l1141381 at Prom 411 IL. ) ... .... S3
17 Failure of Litter Stanchion ............... ........ ,14
1 Te S tUp (ailure at Arrow) ...... 35
19 'Failure of PDC 44 .............................
20 Failure of Aft SlidingL g of Crew So? ................. 7v
RTD-1rDR-63-4230LIST Or ILLUSTRATIONS (CON'D)
F Igure Page
21 Position of Crew Seat Legs in the Most ForwardPoaliton .................................... 38
22 Condition 2B 140 percent (Run No. 2) ForwardFuselage Section at 140 percent ................... 39
23 Condition 313 140 percent DUL -- Failure of Framen555,575.5 and 594 and Rivet Scar In 4eaded WebWhich Attaches to Longeron 114S3356-17 (ViewLoc ing Up) .................................. 40
24 Condition 313 140 percent FUL--Fallure of BeadedWeb Which Attaches to Longeron 114S3356-17(View Looking Aft) ....................... 41
25 Condition 3B 140 percent DUL--Skin Tear andPermanent Buckle at Station 594 (View LookingInboard) .................................... 42
26 Condition 3B 140 percent DUL--Failure of Frame594 (View Looking Up) ........................... 43
27 Condition 311 140 percent DUL--Faillure of Frame,34 (View Looking Forward) ....................... 44
28 Condition 313 140 percent DUL--Fallure of FrameS.34 (View Looking Aft) ........................... 45
29 Condition 3B 140 percent DUL--Failure of RampFairing Aft of FS 482 (View Looking Forward) .............. 46
X() Condition VC--F1allure of Aft Gear at 127 percentDUI .................................... 47
31 ("ondition VC--Pallurv of Drag Llne !141.2029-1at 127 percent [UL ............................ 48
32 Condition VC--F allure of Drag link 1141.2029-1 ......... 49
'3 Wrinkles In New Web Installation .................... 14
iV
RTD-TDR-63-4230
LIST OF SYMBOLS AND ABBREVIATIONS
C.G. center of gravity
MACG most aft center of gravity
MFCG most forward center of gravity
NCG normal center of gravity
Fwd forward
BL buttock line
FS fuselage station
Ult. ultimate
psi pounds per square inch
Ibs/sq ft pounds per square foot
FDTI AF Flight Dynamics Laboratory,Structures Division. Structures TestBranch
DUL design ultimate load
RTV room temperature vulcanizing
WL water line
Nx load factor in the x axis
Ny load factor in the y axis
N7 load fa:tor In the z axis
x angular acceleration Pbout the x axis
y angular acceleration about the y axis
Z angular acceleration about the z axis
vii
'V
iC'i-TiA1 -63-4230
ilvrRoIU(i ICTN
The Vertol CII-47A helicopter was subjected to a complete static test program thatcovered all of the critical flight, landing. take-off, ground handling, and crash conditions.
,%t the completion of the test program two growth testo were conducted. Thu resultsof these tests demonstrated that the CII-47A bisic structure has approxitely 40 per-cent more inhierent strength than was assumed In the original stress analysis.
Several structural failuren and design deficiencies were encountered during the testprogram. These structural deficiencies and the corresponding recommended changes arediscussed In the section on Test Conditions, Dotes of Tests, and Test Results. With thesedeficiencies corrected., the CH-47A helicopter Is structurally capaeblir of withstanding thestatic ultimate loads generated by the conditions outlined in the Appendix.
TE.ST ARTICLE ANT) LOAD APPLICATION METH1ODS
Trhe test article consisted of a complete C1i-47A airframe. A floating test set-up wasusied for all major structural tests. In this procedure, the entire airframe Is tested anonc Integral unit, with the dead weight of the structure and all attached test fixturesrelieved by lead weights suspended from pulleys and attached to the test article. Thiseffectively puts the structure In a zero "g" condition. All test loads were required to heperfectly balanced In translat ion, roll, pitch, and yaw. (See Figures 1 through 3.)
Severasl methods were used in applying the test loads to the fairings and nose enclosure.Neoprene sponge rubber tenision pads and metal-to-metal tension plates bonded to thesiurface with General Electric RTV silicones were used for Introducing tension loads to thesk In (Figure 4). Cobles were attached to each of these loading pointa and were Intercon-nceted to produce particular loading distributions by means of aluminum "whifle treca.-IPbe tension loading on the engine cowling was accomplished by an arrangement of airbladders (Figure 5). Bladders placed in a wooden enclosure and held against the structurewere Lined for compreqsivc air load tests (Figure 6). In all other test conditions notrcquiring air loads, the loads were applied by means of hydraulic rams or atrts. Majorcomponents such as engines, transmissirins, and rotor groups were loaded independent 1vthrough direct attachment fittingn. P'allets vere used for frame and carpo lading In thecenter fuseclaige section (Fi1gure 7). Aluminum tension straps, bonded to the akin with RrVmiliconcot, were used for Introducing loads to the fuselage frames forward of the forwardlr-Ilict! and aft of the aft splice (Figure 8). Treat-load control was accomplished with I disonHlydraiulic Load Maintainces and manual hydraulic control units. The manual units wereused primarily for control of pitch, roll, and yaw In the floating test set-up. A valve wa.inotalled orf the Edison units to rapidly reduce the applied loads to rero in the even of axtrucrurai failure occurring during a test condition.
Mlanutcripm released bvy the author November 1961 for publication an an WrlT) Vechnkicaom )oumenta ry Report.
RTD-TDR-63-4231i
The test article was loaded In 10 percent Increments up to 67 percent design ultimateload (DUL). then was loaded in 5 percent increments up to 100 percent DUL. Deflectiondad strain gage data were recorded from 20 percent through 90 percent DUL in 10 percen:inc rementia.
INSTRUMENTATION
The strain gage installation on the CH-47A helicopter was accomplished by a Jointeffort of FDTT and Vertol. Approximately 400 Bud Company C12-141 and C12-121 gages(foil, epoxy backed gages which are temperature-compensted for aluminum) were in-stalled. Critical strain gage and defle.-tion transducer locations were selected prior toeach test condition. Strains were recorded at selected incremental percentages of DCLwhile deflections were recorded continuously or at the same increment as strain. Criti-cal applied loads were monitored with calibrated load cells to insure accurate loading, toeliminate overloads, and to determine the actual percent of load at which a failure oc-curred. Nosker Engineering Products Strain Indicators (Model 2A), which had been modi-fled so as to be compatible with a Gilmore Industries Graphical Recorder-Plotter (Model114). were used to monitor strains. This system was capable of automatically plotting 144data channels at a rate of one sample per channel per second. The total system accuracy(from the strain gapge to the recording instrument) was *5 percent of full scale. "fuliscale" being the full scale range used or 5000 micro inches of strain. Deflections weremonitored by a system comprised of Research, Incorporated Displacement Transducers(Models 4040 and 4046), Research, Incorporated Transducer Control Cabinet (Model 4095).and a Research, Incorporated Recorder-Controller (Model 4060). The total system ac-curacy was ±I percent of full scale.
DESCRIPTION OF STRUCIRE
The Vertol CII-47A helicopter has twin rotors arranged in tandem fashion. The frontroor is supported by the front pylon structure, which in located directly above and aft ofthe fuselage cockpit: and the aft rotor, supported by the aft pylon, is located at the extremvaft end of the fuselage. The helicopter Is equipped w.!th a fixed quadricycle type alightinggear. consisting of a fixed vertical type for the forward gear and an aft landing pear havingfull castering capability.
The forward transmission and rotor group are suspended by four fittings, one on eachside of a box-type super structure extending from the crown of the ship.
NOTE: A detailed description of recording Instruments, transducer characteristics.methods of Installation, type of output information, and transducer location. on theCII-47A are available to contractors through the Air Force Flight lDvnamiclaboratory, FPiTT.
RTD-TDR-b3-4230
The primary structure of the fuselage Is a semi-monocoque construction with sup-plementary structural members provided In the vicinity of open sections and at thoselocations where large concentrated loads are Introduced. The smi-monocoque structureconsists primarily of the following: (1) longitudinal members to provide bending restraintin the fuselage; (2) metal skin to carry the fuselage shear and torsion; and (3) transverseframes and bulkheads to provide stiffness In the skin structure and to maintain the fuselagecross section during flexure.
The fuselage skins consist of 2024 aluminum alloy alclad sheet for the top and side skins.while 7075 aluminum alloy alclad sheet is used on the bottom portion.
The aft pylon consists of an upper and lower section. The primary structure In theupper portion is a closed box form. bounded on top by the thrnit deck, on the botom bythe torque deck (WL 72), at the ends by bulkheads at Station 482 and 594, and along thesides by the skins. The primary structure below WL 72 consists of two open end frames(Stations 440 and 482) and the crown and side skins.
TEST CONDITIONS, DATES OF TESTS, AND TEST RESULTS
)uring 81 tests run, the C}!-47A was subjected to 71 different test conditions. Thesetest conditions are explained in Tables I through 6 (Appendix) and include conditions forflight, landing, air load, flight controls, components, and ground handling. In the listingwhich follows, some test numbers are preceded by an asterisk. Explanations of tests somarked follow the listing.
Ultimate'rests CFI-47A Test Condition Test Date Load Supported (%)
I Flight control system cond. No. 7 7 June 1962 100
2 Flight control system cond. No. 8 7 June 1962 100
3 Flight control system cond. NW. 1 8 June 1962 100
4 Flight control system cond. No. 2 8 June 1962 100
5 Flight control system cond. No. 3 3 June 1962 100
6 Flight control system cond. No. 4 11 June 1962 100
7 Flight control system cond. No. 5 11 June 1962 100
8 Flight control system. cond. No. 6 12 June 1962 100
Q Flight control system cond. No. 6 12 June 1962 100
10 Flight control system cond. No. 11 20 June 1962 94
I l Flight control system cond. No. 10 20 June 1962 00
i3
RTD-TDR-63-4230
UltlmsteTests CII-47A Teat Condition Tem Date Lmd Suipported (€
12 Flight control system cond. No. 12 21 June 1962 100
"13 Flight control system cond. No. 11 28 June IQ62 85
14 Escape panel airload test 6 July 1962 100
15 Flight control system cond. No. 14 20 July 1962 100
16 Flight control system cond. No. 15 23 July 1962 100
17 Flight control system cond. No. 13 23 July 1962 100
18 Load-defection test No. 1 11 September 1962 N/A
11) Cockpit enclosure unsymmetricalflight cond. 28 September 1962 100
20 Cockpit enclosure symmetricalflight cond. 4 October 1962 100
21 Main cabin door air load test 11 October 1962 100
22 Leading edge fairing of aft pylon 18 October 1962 100
23 Forward pylon fairing 30 October 1962 100
24 3B MACG 3.3000 pounds gr3asweight 4 February 1963 100
25 3B MACG 27100 pounds grossweight 13 February 1963 100
26 6B MACG 27106 pounds grossweight I March 1963 100
27 713 MCG 33000 pounds grossweight 5 March 1963 100
28 23 MFCG 33000 pounds grossweight II March 1963 100
29 2B MFCG 27100 pounds gross
weight 12 March 1963 100
*30 External cargo tow hook 3 April 1963 67
31 Exters! carg hook 16 April 10
4
RTD-TI)R-63-4230
UltimateTents CFl-47A Test Condition ITest lae Load Suppored()
32 Main landing gear cond. VC(50-50 distribution) 30 April 1963 I0)
13 Main landing gi*Ar cond. VC
. (60-0 distribution) I May 1963 100
034 Main landing gear coed. VIB 3 May 1963 97
35 Fngine cowling airload test 3 May 1963 100
:6 Aft gear towing condition 8 May 1963 100
37 Forward gear landing cond. IVD(60-0 distribution) 17 May 1963 100
:8 Forward gear landing cond. IVD
(.'-50 distribution) 20 May 1963 100
.NL Forward gear towing condion 23 May 1963 100
40 1.xternal cargo hook
(track beam assembly) 28 May 1963 100
41 Rescue hoit 3 June 1963 100
42 isting condition 5 June 1963 92
43 FWI) transmission and rotorgroup 8G down crash cond. I I Junoe 1963 100
44 l'roop commander seat aide load 13 June 1963 100
45 Crew seats - SG down crash cond. 14 June 1963 100
46 Troop commander seat (back load) 14 June 1963 100
47 Crew seats - SC side crash cond. 17 June 1963 99
48 Troop commanders mot max.down load 17 June 1963 100
49 Troop commanders seat lap beltIrad 24 June 1963 125
Ni "rroop seat lap belt load 24 June 1963 135
Sl lone man seat lap belt load 25 June 1963 100
- 1 1.
RTD-TDR-63-4230
UltimateTests CH-47A Teat Condition Test Date Load Suppomed ()
52 5000 pound carso tie-down sup-ports Test No. 1 26 June 1963 100
53 5000 priund cargo tie-down sup-ports Test No. 2 26 June 1963 100
54 10000 pound cargo tie-down sup-ports Test No. 1 26 June 1963 100
55 10000 pound cargo tie-down sup-port.'Test No. 2 26 June 1963 100
56 Littcr 4.50 down crash cond. 1 July 1963 100
57 LIter 1.5G sloe crash cond. 2 July 1963 100
58 Cargo ramp and support struc-ture 2 July 1963 100
059 Litter 8( forward crash cond. 2 July 1963 62
60 Cargo W:adling 8 July 1963 100
61 Engine 8G Fwd crash condition 12 July 1963 100
62 Engine 80 down crash cond. 12 July 1963 100
63 Engine G side crash condition 15 July 1963 100
64 Main combining transmissionBG Fwd crash 15 July 1963 100
65 Main combining trasmmlaion8G down crash 15 July 1963 100
66 Main combiting transmiasion8g side crash 16 July 1963 100
067 Hoisting condition 17 July 1963 100
69 Aft transmission and rotor groupRG down crash 18 July 1963 100
69 Aft trsnsmlslon and rotor groupAG Fwd crash 22 July 1963 100
7') Fwd transmission and rotorgrnup G Fwd crash 23 July 1963 100
6
RTD-fl)R-63-4230
IltimateTests CH-47A Test Condition Test Date Load Supported ( )
071 FIIht control system cond. 11 26 July 1963 100
072 Crew sests-aG Fwd crash cond. 26 July 1963 80
073 Crew seta-OG Fwd crash cond. 29 July 1963 125
074 2B'MFCG 33000 pounds grossweight 6 August 1963 140
075 30 MACC 33000 pounds grossweight 12 August 1963 140
076 Litter 8( Fwd crash cnnd. 14 August 1963 72
$77 External cargo tow hook 20 August 1963 100
,78 Max. floor load 23 August 1963 100
079 Troop seat 8( side load 26 August 1963 80
080 Troop seat SC down load 27 August 1963 90
081 Landing cond. VC (new shearweb Installed) 27 September 1963 127
Tests I0. 13., and 71
During the test of flight control system condition No. II, a failure occurred at 94 per-cent DUL. The 114CI013-8 push-pull rod failed In compression. The boost pressure atthe time of failure was 1500 pi. A similar rod was tested with a boost pressure of 2250psi. The rod failed In compression at 85 percent DUL. A new rod was dealned with agreater wall thickness and retested. The system atisfactorilv supported 100 percent DULwith the new rod (114C1013-33). (Figures 11 and 12.)
rtmts .0 and 77
During the test of the external cargo hook, several strain page readings were abovethe allowable indicated by Vertol at 67 percent DUL. Further analysis indicated that theoriginal allowable strain was too low, therefore the external cargo hook was retested andsatisfactorily supported 100 percent DUL without failure.
r'eat 34
The aft landing par ohear box falled at 97 percent DL in condition VI (Figures 11and 14). This was the third major test on this gear and there were shear wrinkles priorto this test. The structure is considered satisfactory In view of damalge Incurred duringthe previous tests.
RT)-TDR-63-4230
Tests 42 and 67
During the hoisting condition, a failure occurred in the aft crown boom Instaliatie at92 percent DUI. (Figures 15 and 16). The main web of the aft crown bmam installation(Part No. 114-S-3351) was strenphened and the condition was retested to 100 percent IXJI..
Tests 59 and 76
During the litter SC forward crash condition, a failure occurred In the litter stanchionat 62 percent DXUL (Figure 17). During retest, with the lower plate of the stanchion Increasedin thickness, a failure occurred in the litter strap floor attachment fitting at 72 percentIAJL (Fligures 18 and 19). Further tests are necessary to prove the litter Installatkt ustim-factory.
'rests 72 and 73
At 80 percent DUL, a failure occurred in the sliding leg of the crew seat during the BIGFwd ciash condition (Figure 20). This failure was due to the position of the seat, In thatit allows only partial contact of the shoulder of the meot leg with the seat track flangeswhen the seat is in the most forward position (Figure 21). The moat was placed one inchaft of its most forward position and satisfactorily retested to 125 percent DUL. It wasrecommended that the seat adjustment be relocated to prevent the seat from being placedin the present extreme forward position.
Fest 74
Condition 2B MFCG 33000 pounds gross weight was tested to 140 percen DUL for a growthpotential indication. This condition Is critical for the forward pylon and splice, ans Indicatedin the Appendix. The structure supported 140 percent DIJl after the forward transmissionsupport structure (Part No. 1 14St IRA-193) was restrained from buckling at 12S percentDOL. Only minor damage was noted (Figure 22).
'Feist 75
Condition 313 MACC 33(X) pounds groan weight was tested for growth potential. Thiscondition in critical for the aft pylon, as indicated in the Appendix. The following failuresoccurred at 140 percent Ill,:
(a) Failure of frames 555, 575.5, and 594 (Figure 23):(b) Failure of longeron 114S33.56-17 (Figure 24);(c) Skin tear and permanent buckle at P.S. 594 (Figures 21 and 26):(d) Failure of frame 534 (Figures 27 and 28):(e) Fallure of ramp fairing aft of F.S. 482 (Figure 29).
rests 79 and 80During the troop seat R( side load, a failure occurred in the leg strap of the seat at IO0
percent IXJL. The hack panel of the troop seat was aga:nt the foselage at 40 percent.
The seat back straps failed at 90 percent IYJL. during the troop met AG down crashcondition. The troop seat back-up st ructure could not be tested to the SC load conditionabecause the existing troop seats are designed to a lower "r" limit than the hack-upstructure. In testing the hck-up structure, a troop weat must be used so an to dixtribute
+S
RTD-TIYA -63-4230
the load corretly throughout the back-up structure.
Teat 81
The aft landing gear condition Vc was retested with a new shear web Installed. The draglink on the main gear forging (Part No. 114L 2029-1) failed at 127 percent lxUI (Pigures30. 31. and 32). The new web Installation was permanently wrinkled (ligure 33).
CONCLUSIONS ANt) RECOMMENDATIONS
Based on the results and obervations obtained from the Ctl-47A Static Test Program.the following conclusions and recommendations are presented.
1. As a result of static test failures or deflencies; enroumered during the testprogram, FLIT recommends that the changes Indicatod In Tots 10, 42, 59, and 72(preceding section) be Incorporated on all CII-47A service helicopters.
2. In the event that a troop mt Is designed to take a greater load than the existIngtroop seat, it is recommended that the back-up structue be tested to DUI. with this seat(reference Test 79. preceding section).
3. The results of the growth testing Indicate that the baesic fuselae structureof the CII-47A has approximately 40 percent more Inherent strength than was assumed Inthe original design strew analysis. However, the forward transmislion support structurerequired additional stiffening. The reader should not@ that the landing lear and other com-ponents have not been tested for overloads.
4. In general, the measured stresses recorded during the tests were lew then thosepredicted by stress analysis.
5. With the Incorporation of all the structural changes recommended In this report,the CII-47A Iielicopters are capable of supporting the static ultimate loads for all of theconditions listed in the Appendix.
6. Fromn test data and observations, the areas with the lowest margins of safety,althouOh they supported 100 percent DUL are,: (a) The upper forward corner of the maincabin door: (b) and the aft landing gear back-up structure.
iI
CII -47A STRLJCTURA L TES-T ('ONDrnON
RTD-T)R -63-4230
FAISI I
CLII-7 P"aLlTed flla * ArgIIJaw
[cot COndicbmui (Lba cti (. . Nx NY N ~ x v £ riaArcaf
291 .44 141 M( 1- 4.12 -.21 Fuid pyl~i mAI P.wdxyettettric.sl slt picet critical inand pull-tin.t n'WW wwfllcal "mrinif
topt pitchinm
:~dea II- i~tm 3NI0N Mm ..M I.. ~.Ii 1."t .291 Aft pyluui CriticalOyelt"mri'rkI divv In verfiaal tw nurc41vwery frutit
'iwnriWtchiW
GordlIIe 311- 27,11110 NIAUO .40 0t 1J.6I -.171 1.4h .2401. R'Wir LI."onericral divtrcciefry frtpm
itw. ri~ht
cin~diuilsn 'M1 14hWu NIlo *.-X t4 NUSI41 ii i i aurt i'nd
mtal'u lE IV,
RII )- t ilt -t l-42Ml
TABLi+ 2
CII-47A LANDING COM)ITIONS TISTI:I)
Gros Wt.Test Conditions (Lb) C.G. Critical Arelsi
Conltion VC side ob- 27.100 MACG Aft gear and hack-up "truc-struction 50-50 distril- ture (shear box)but ion
Condition VC side oh- 27,100 MACG Aft par axlesuruction fiat tire con-dition 60-0 distribu-tion
Condition Vs aft oh- 27,100 MACCs Aft par and hack-up ntruc-atruction ture (shear hox)
Condition IVI) side 27,100 MI(CG Pwd gear and hack-up mruc-obstruction 50-50 turedistribution
Condition IVI) aide 27,1(X) MFCG Fwd par axleobstruction flat tirecondition 60-0 dis-tribution
TABLE 3AIR LOAD CONDITIONSTEI)
Velocity Yaw of AttackTecat Condition (K not a) (D~egrees) (Degrees) Critical Area
metrical flight conditionIAnlCockpit cnciosure--syn- 153 I1 -1.7 1 Cockpit cnclomsure(0c kplt snchoeurc--un- I Ii) !1% -I.7 C (ockpit enlcloaur
ynlmctriz.'ai flight con- Idtl Ion
Forward pylon fairing 14S 20 -1.7 Work platform ll4.l'i
Leading edge fairing i%' I% -2.I Ieading edge dmirtoof aft pylon
[scpe panel P I -1.7 l'anel mouldin!g
Main cabin diior IRI -1.7 I )NW moulinVIUMilg Mki.itd
ment a
Engine cowling lil i I1 igIn in ttln ,.Pr
14
RTt)-TI)*t, -6142.30
rAII.L. 4:LIGIrr CONTROLS TE.ST rONDTOM
1 3i
Test Load Application I Control Ifhimatc I.4=41
Condition System ! ffort Point ositaIon (1.1w)
1 LongI- Single pilot I'lot's hand grip Fwdtudinal
2 Longl- Single pilot PIllot's hand grip Aft 34K)
tudlnal
3 I.ateral Single pilot Pilot's hand grip Left I.,
4 Iateral Single pilot Pilot's hand grip RIght 5IA
5 Collcc- Single pilot i'llot's hand grip Down 22S
6 Collec- Single pilot Pilot's hand grip tip 22S
I)Iree - Single pilot I'lot'm right pedal Right 491kinil i pedal aft
A I)rec- Single pilot i'ilot's left pedal Left 4S1
t (WPl pedal I aft;
Longl- I)uaI pIlot Sticks Fwd t 225rudinal plus boost cacti
I10 l.ateral lDual pIlot Sticks Left 112.5ph11 boost each
I I Collce- l)ual pilot Sticks Up 169tIve plua boost each
12 I)irec- Dual pilot RIght pedal RIght ,37.Silonal plus boost pedals aftl each
1I Rotor N/A wd rotorhead AIMind ft. .ft rotorhead AIN)
14 Roltor N PA Fwd rotorhead 2MO
and fit. Aft rotnrhcidcon rols
I Rotor N,1A Fwd rotorhcadand fit. Aft rotorhead.
"" cont rolsL .... ......... ... ..... ...... ....
a 1
- I
WI1)-lIDK-63-423-1
TABI.I ,sCON1,IONINT TIEST CONDITIONS
rest Condition- Ultimate .o d
inginc W, Fwd crash conditzi M6M per eninc
Lnglne RG down crash condition 6656 per engine
Lngfne 86 silde crash conditlm 6656 per engine
Forward transmission and rotor group 86 down crash 22139condition
F'orwar transmission and rotor group 86 Fwd crash 2213)condition
Main combining transmission 8G side crash condition 2574
Main combining transmission 8G down crash condition 2574
_ Main combining transmission 8G Fwd crash condition 2574
Aft rotor group and transmission 86 down crash 25106cold ition
Aft rotor group and transmission 8G Fwd crash 25106condition
Crew 'teat 8(3 Fwd crash condition 1920
Crew scat 8C side crash condition 1920
Crew scat 86 down crash condition 1921)
I'rrp t4eat supporting sructure (lap belt load) 2030 per seat
Troop seat 86 side lxd 2080 per seat
Trxp seat 1W down load 20W0 per xs
(e man seat lap helt load 20.10
f'roop commander neat (maximum down load) 2000
fr(xp commnder -way side l(ad 22S
rr~op commander meat (thack load) 001
1 r(xp eommasnder nreat lap belt had IMOO1
Ito
iTI)- ll)i{-63 42011
I'IMILlS (CONr'.)
l'+st Condition Ultimate Load
I.Itter inttallation supports 1.50 side load 375 per litter
.irter installation supports 4., down load 1125 per litter
l.itier installation supports 80 Fwd load 20(0 per litter
Re. (ue hoist 180
( Cargo handling 4M)O
I'xternal cargo hook (back-up structure) 48,000
External cargo hook (track hean assembly) 48,000
I *xternal tow hook 69,212
Cargo ramp 9000
Hoisting condition 60,000
(arg tic-down supports (10.000 pound fitting) 10.000 (45 deg. aft),
Cargo tie-down supports (10.000 pound fitting) 10.000 (3 deg. aft)
Cargo tic-down supports (5r(X)0 pound fitting) 5000 (45 deg. aft,)
('argo tic-down supports (50) 0 pound fitting) 5,0X) (3 deg. cift)
Max. floor loading :300 pound per square foot{ 210W IISF7G
GROUND IANI)l.ING coN)mIT IONs rlsrl)
l'egi Condition row Load Vertical Ld(Lbs UI.Tr iII' I'l U',
lorward Gear 'rowing CokItion 10.422
Aft.car Towing Condition 8,44 ,64
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