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FOR FiRH, T AIiOZ USARTL-TR-78-22
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CH-47 CRASH TEST (T-40) STRUCTURAL, CARGO RESTRAINT,AND AIRCREW INFLATABLE RESTRAINT EXPERIMENTS
L. Burrows, R. Lane, J. McEIhenney
D C
April 1978 DDC
Reproduced From dforpublic release;
Best Available Copy
Prepared
APPLUED TECHNOLOGY LABORATORYU. S. ARMY RESEARCH AND TECHNOLOGY LABORATORIES (AVRADCOM)Fort Eustis, Va. 23604
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PREFACE
The Project Engineer for the fuli-scale crash test Jescribed herein was Mi. LeRoy Burrows,Applied Technology Laboratory, U. S. Army Research and Technology Laboratories(AVRADCOM), Fort Eustis, Virginia. The cargo restraint experiment was conducted byMr. Richard Lane, Applied Technology Laboratory, USARTL (AVRADCOM). Mr. JamesMcElhenney, Naval Air Development Center, Warminster, Pennsylvania, conducted theinflatable restraint experiment.
The authors extend their gratitude to the following Individuals for their contributions tothe efforts documented herein:
IT, C. Castle, NASA-Langley Research CenterG. Sinyley, IlI, Applied Technology Laboratory. USARTL (AVRADCOM)R. Bott, Applied Technology Laboratory, USARTL (AVRADCOM)L. Domzalski, Naval Air Development Center
Also, the authors are appreciative of the following organizations for the support specified:
1. NASA-Langley Research Center for facility support, conducting the tests,instrumentation support, data recording, data reduction, external photography,and test vehicle disposal.
2. U. S. Army Transportation School Aviation Maintenance Training Department,Fort Eustis, Virginia, for conducting the weight and balance analysis, for pro-viding CH-47 mechanic support, and for the loan of aircraft equipment and amaintenance stand.
3. 355th Helicopter Company (HH) (7th Trans Group) for transporting the testaircraft from Fort Eustis to the Langley test site.
4. Fort Eustis Maintenance Division for interior and exterior painting of thetest vehicle and loan of aircraft jacks.
5. New Cumberland Army Depot for furnishing CH-47A fuel cells, tires, andinertia reels.
• LT•lng4 1/ • 'I071F~. _I_11 -.,__..... ..... ......
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4 TABLE OF CONTENTS
Page
PREFACE ........................................ ............... 3
LIST OF ILLUSTRATIONS .................................... .. .......... 6
LIST OF TABLES ..............................................................- ........ 9
INTRODUCTION ...................................-.............................. 10
TEST FACILITY...................................-....................................... 11
4TEST SETUP........................................................................ . ..... 15i
Aircraft .................................................................. ....... ~ 15Instrumentation ....................................... ................ 16Photographic Coverage .............................................................. 19
TEST DESCRIPTION.................... ......................... .... . ................. 23
Procedure..................................................... . .. ..... 23Impact ........................................................................... 24Data Acquisition ............................................. ~... 28Date Reduction and Presentation ................ . ........................ 28
*EXPERIMENTS ........................... .-..................... ..... ...... 29
Structural ............................. ......................... ... ..... ...... 29
Purpose ................................................ ..... .......... 29..Background ................................................................. 29Description ................................................. .................. 29
*Results ................................. ................................... 30Conclusions .......................................... ............. 36
Cargo Restraint ......................................................... 6 ............. 7
Purpose ........................... ...-............................... .... 57Description ..................................... .......................... 57Results ............................................. ... ..... ........ . 59
*Post-Crash Laboratory Test .................................................... 63Conclusions ....................................................... ........... 63
Aircrrwv Inflatable Restraint .................................. ...................... 66
Purpose........................................................................... 66Background ..................................................................... 66Description ........................................... ....................... 68Rqsultu ....................................................................... 69Conciua~cns.......................... ........................................ 80
APPENDIX A -Instrumentation ......................................................... 81fT.
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LIST (;F !-LLLJ3FRTR 1\7lIS
F Page
1 1iL;,.rn2ut yn.1ts~ ch Facility................................ 12
2 'Systcn) uý,i-d in fu!1 c;~ ri:'n silmulation test............................. 12
3 Lift~ir,_ III Jid2&;n ................................................. 13
4 Cber!: yei......................................................... 14
5 Tc~t -p;.chM1ýn ................................................................... 15
0 CH-17 T-40 ................. .....)............................... 17
7 C11-47 T-",O cz~r-,ra and lo,~t cations .... ................................ 21
0 m~c .................................................... 25
9 Irnp:ct Squ~tce ~t.J viwlv...... ...................................... 26
10 o~t-craih vi.-v riht front................................................... 27
11 Vost-crch vivw loft r.-c-r...................................................... 27
12 FucIl ccil rupt;,'ro and \v;Mtcr dissipation.................................... 30
13 Po-,t-cr,,ih vicwi of fujil t-od and ce'll........................................ 31
Structurcl f;ailuic at 2A~c .......:..................................... 32
153 Intcrior ficor d:;f.orm-,ticn at station 2,10O...................... .......... 311
....... ............................................................ 33
17 -ý- pnin f"-ihura ............ .................................................... 33
12J V-r~i.-on in v aIrtin nd czabin heirht reduction with........................................................................ 3
................................................................... 37
......................................................................... 3
C!';: fl::,r ;ýnd 1 -rr attachm'2nt vcrtical decelerations (rawcrr 1C I"! ................................................................ 41
Figure W
22 Cabin floor and landing gear attachment vertical decelerations(100 Hz filtered) ................................................................................... 43
23 Cabin floor and landing gear attachment vertical decelerations(velocity Integration) ........................................................................... 45
24 Cabin floor and landing gear attachment vertical decelerations(displacement integration) ................................. 47
25 Cabin height reduction ......................................................................... 49
26 Cabin floor longitudinal decelerations and fuel cell pressure .............. 51
27 Pilot station decelerations ................... ................................................... 6328 Copilot station dece~erations ............................... 55
29 Rear pallet load looking forward, after impact ............................... 58
30 Cargo load attenuator ......................................................................... 58
31 Rear trailer looking forward from starboard side, after impact ....... 60
32 Rear trailer looking forward, after Impact . ............ 60
33 hear trailer, aft restraints with energy absorbers, before impactfstation 340) .................................. 61
34 Rear trailer, aft restraints, after impact (station 340) .............. 61
35 Forward trailer, forward and lateral restraints, after impact(station 240) .......................................... 62
36 Forward tra'ler just aft of bulkhead ......................... 62
37 Post-crash load attenuator laboratory tests ......................................... 64
4 38 Partially extended energy attenuator showing interference of plattenwire with tension adjustment wheel on MB-1 chain restraint ....... 6.5
/ 39 The Inflatable re-.raint system ........................................................... 67
40 Comparison ol Inflatable restraint in stowed position and unfurled 67
S41 Thiokol Chemlical Corporation's pyrotechnic Inflator .......................... 68
42 CH-47A test vehicle (post-test) ............................. 69
. |7
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43 Post-tost cockpit showing crash damage and dummy position......... 70
44 Pilot restraint shoulder loads versus time ................................. 71
*45 Copilot restraint shoulder loads versus time ............................. 72
46 Pilot restraint hip loads versus time....................................... 73
47 Copilot restraint hip loads vcrsus time.................................... 74
O48 Pilot and copilot resultant chest acceleration versus timr ............. 75
49 Pilot horizontal chest acceleration and bladder pressure versus time . 76
so Aircraft horizontal acceleration and bladder pressure versus time .... 77
51 Aircraft horizontal acceleration and velocity versus time.............. 8
52 Aircraft vertical acceleration versus time ...................... 79
A-11 Army recording system...................................................... 93
AA-2 NASA recording system ..................................................... 88
A-3 Deflection tube Installation ................................................. 89
A-4 On-board camera/light system .............................................. 97
A-5 Arrny junction box........................................................... 99
A-6 NASA junction box.......................................................... 99
A-7 Instrumentation cabling techniques ........................................ 100
8
LIST OF TABLES
Table
I Weight and balance components ........................................ 16
2 Crash test impact parameters ........................................ 23
3 95th percentile potentially survivable accident design pulse ........... 24
4 Summary of drop tautdata............................................. .70
A-1 Recorder N functions and calibration data ............................. 84
A-2 Recorder 0 functions and calibration data ........................ ... 85
A-3 Recorder P functions and calibration data ............................ 86
CA-4 Recorder A functions and ca~ibration data ............................ 90
A-5 Recorder B functions and calibration data .............................. 91
A-6 Recorder C functions and calibration data ......... ................ 92
A -7 Recorder D functions and calibration data ............................. 93
A-8 Recorder E functions and calibration data .............................. 94
A-9 Recorder F functions and calibration data ........................... 95
INTRODUCTION
As part of its crashworthiness R&D program, the Applied Technology Laboratory hasbeen involved in 39 full-scale aircraft crash tests during the past 16 years.
On 4 August 1976. the 40th test (termed T40) was conducted at the NASA-LangleyResearch Certer's Impact Dynamics Research Facility, Hampton, Virginia, The testvehicle was on Army CH-47A troop/cargo helicopter. The objective of the test was to""lAtain data that would contribute to the development of helicopter crashworthiness fee-S. turins and design criteria, thus enhancxing occupant and cargqo survivability. This was4accomplished by three instrumented onboard experiments:
* Structural
* Cargo Restraint
0 * Aircrew Inflatable Restraint
The program was a joint Army/NASA/Navy effort. Tests were conducted by personnelfrom the Applied Technology Laboratory. U. S. Army' Research and Technology Labora-tories (AVRADCOM); NASA-Langley Research Center; and the Naval Air DevelopmentCenter. NASA provided the test facility, conducted the test, installed some instrumenta-tion, and reduced !he test data in accordance with requirements specified by the Army.The Army was responsible for the experiments, the vehicle readiness, and the overallproject management. The Navy installed the inflatable restraint system.
This report documents the test facility, setup, description, and experiment results.
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Figure 1. Langley Impact Dynamics Research Facility.
Pivot point plaltoris Plbc al
Camera iocation (typical)
Figure 2. System used in full scale crash simulation test.
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TEST SETUP
AIRCRAFT
The test specimen was a CH-47A helicopter, S/N 61-2418B, which had been rotired fromservice. It was equipped with both engines, all transmissions, and stub blades. The air-craft was ;nitially prepared for testing at the Applied Technology Laboratory (ATL), FortEustis, This preparation included cleaning, interior stripping, installation of new fuel cells.plexiglass removal, interior and exterior painting, and preliminaty weight and balancecalculations.
* Mounts and fixtures for the on-board cameras and flashbulbs were designed and fabricatedby ATL personnel, as were the pellets, trailer loads, instrumentation mounts, deflectiontubes, and many other items supporting the experiments.
The interior was painted white and the exterior yellow with black lines, indicating thelocation of bulkheads and longerons of interest. An exterior view of the test specimen isshown in Figure 5. The test aircraft was transported from Fort Eustis to the Langley testsite, where an ATL team, assisted by NASA and Naval Air Development Center (NADC)personnel, installed the experiment hardware, instrumentation, cameras, lights, and fixturesaboard the helicopter. Standard rotor blades could not be used as they were not compati-ble with the hoist rigging concept. Stub blades were installed to provide some representa-tion of the rotor mass effects on the main transmissions.
•. •Figure S. Test specimen.
- -15
A final weight and balance calculation yielded a C.(i. position at station 323.7 and anaircraft gross weight oi 25,010 pounds, which is close to the normal desigrn gross weightof 25,500 pounds for a CH-47A helicopter. A summrry of the weight components isgiven in Table 1. The fuel cells were filled with wier to the weight consistent with afull load. Ballast was placed in the closets at edtion 107.5 to aid in obtaining a C.G.that was within an ;nch of that stated in the CH.4)A model specification as the normalC.G. location.
TABLE 1. WEIGHT AND BALANCE COMPONENTS
Item Weight (Ib) Station Moment
Sasle aIrcraft 13,488 354.5 4,781,46
Fuel cells 60 316.2 19.006
Water In fuel cils 3,811 316.8 1.207.325
Oummile ltwo) 3W 71.1 29.289
Aft trailer and restraInts 1.426 369 626.194
Forvwd trailer and restraints 1,376 166 227.370
Aft pallet and restraints 834 462 385,306
Forward pallet and restraints 814 280 227960
Baseline cargo package 179 326 58.354
Datterles and mounts 152 300 45,600
Forward stub blades 384 66.7 33.293
Aft stub bWKc'e 384 553.7 212.621
Cameras and ftixth.e: No. . a 5 134 370 49.110
266 119 196 23.324
4 36 390 21.450
1 72 133 9.576
Ballast:
Rilght-ht•n heat• cloeet 517 107.5 36,576
Left-han! closet 617 107.5 66.576
Mb•lecellnus Fixtures 296 428.5 12.833
TOTAL 25,010 323.7 8,095,737
Final test preparation activities Included completion of instrumentation installation andconnection to umbilical cables, check of each data channel on an oscilloscope, loadingthe cameras, and Installation of the pyrotechnics.
INSTRUMENTATION
One hundred twenty-eight instruments were used to record data during the test. Theseconsisted of 64 accelerometers; 42 strain gages; 10 load cells; 3 pressure transducers;Doppler radar; and 8 deflection tubes, two of which were rigged for manual measurementof the maximum structural deflection of the cabin ceiling. Figure 6 shows the general
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location and the type of each on-board -instrument and its recorded channel numberdesignation.
Appendix A provides a detailed description of the instrumentation used In this test,and describes the instrumentation mounting, calibration, cabling, circuitry, andand tecording procedures. Because of the importance of obtaining valid crash ac-celeratirin data in the vertical direction, redundant accelerometers for this axis wereinstalled at many locations.
Each experiment section of this report addresses instrumentation supporting that specificexperiment. There was a concerted effort to place instrumentation such that the datawould be applicable to more than one experiment.
PHOTPGRAPHIC COVERAGE
Color motion picture coverage was provided for each of the on-board experiments. NinePhotosonic high-speed 16mm motion picture cameras were installed in the test aircraft,as shown in Figure 7. These cameras were chosen because they are compact, lightweight,and hoave the ability to withstand high impact shocks. Seven cameras were placed in the
* cargo compartment and two were mounted in the cockpit. A protective covering madefrom heavy gauge aluminum was constructed and used to enclose each camera. Theywere then mcunted on a specially designed shock-absorbing mounting plate to reducecamera motion and impact shock. Due to limited camera placement locations inside theaircraft, a 13mm wide-angle lens was used for its wide field of view and its ability to
* obtain visual data at close quarters. Experiment observation requirements dictated thatas much of the tie-downs and cargo be phiotographed as possible.
A 28 VDC aircraft-type NICAD battery placed on-board was used to power the camerasand their control circuitry. The camera control system was activated by an aircraftrelease signal from the drop control room, which allowed camera turn-on and a "run-up"of approximately 1.25 seconds before the light intensity became sufficient for adequateexposure. All cameras, once on, continued to run through their total film supply and
* then automatically shut off at approximately 14 seconds after release. The cockpitcameras were attached to ceiling structural members to minimize camera movement duringthe impact sequence.
Because of the high impact forces to be encountered, Sylvania FF-33 special-purposeflash lamps were selected for interior lighting. Previous impact tests had proven that theselamps were suitable for this type of application. The lamps were installed in 7-inch
* polished reflectors with a wire retaining screen and located as shown in Figure 7. Theinterior of the aircraft was painted white so that reflected light would increase the lightlevel inside the aircraft. A second NICAD battery was installed to power both banks of
* flash bulbs. Bank 1 consisted of 12 flash lamps, while Bank 2 contained 14. The con-trol system was activated by a rear-mounted aircrbft lanyard switch that was tripped uponaircraft release and subsequent movement. Light Bank 1 was designed to activate approxi-mately 1 second after aircraft release and Bank 2 approximately 15 seconds after release.
19
The exterior photo~h vP1'V SIM prvde V qkm.CO ri; ncucdec;thilZ
grudcneTas, twyo of which V4ere on II~c ý,ntri ovcthcýA, w&nd four MCnnsim
4 gro~ind came~~ ~ttio~cd to cover ih viJM (elt, fore, and aft i~icws o ~c ~ v i tr
in %dditioni. NASA provide'd vidoo taPe cQVPTýT fOr irnmý-diite O-tA~p~V
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TEST DESCRIPTION
PROCEDURE
The test specimen was lifted by the two swing cables and pullhack cable winches untilthe longitud:nal and vertical center of gravity was located in the proper impact position,as shown in Figure 5. The swing cables were locked in this position, and the testspecimen was raised by the pullback cable until the center of gravity of the te:t speci-men waes at the desired height of 54.14 feet (Figure 3). This height was calculated toprovide a flight path that would res.lIt in the planned impact conditions given in Table2. The pendulum method of impact testing was ustd because a combined horizontaland vertical velocity resultant impact vector was desired and crash impact conditnnpredictability was important.
TABLE 2. CRASH TEST IMPACT PARAMETERS
Planned Actual
Pitch angle. deg -5 -8.7
Yew ange,, dog 0 0
&._ angle, dg 0 0
Vertical velocity. f/ise 42 43.5
Horizontal velocity, ftlwc 27.1 28.3
ResuItant, velocity. ft/we 50 51.9
The planned crash conditions for T-40 were selected because they would cause a largeplastic deformation of the primary airframe structure and would permit in assessmentof the airframe failure modes in a sever, but survivable crash impact. The nose-downcrash attitude was selected because the previous CH-4C crash test (T-39) had a 10-degree nose-up attitude impact, ard the attainment oi data for this range of impactco'nditions would enhance its usage and empirical 4orrelation potential, especially in thedevelopment of the structural crash simulation computer model, KRASH.
A 50-ft/sec resultant impact velocity vector was selected for both T-39 and T-40. Thisis representative of the 95th percentile potentially survivable crash p.|lse as defined inMIL-STD-1290(AV) and given in Table 3.
23
...
"TABLE 3. 95TH PERCENTILE POTENTIALLY SURVIVABLEACCIDENT DESIGN PULSES
.Velocity change
Direction along aircraft axcis (ftlchil
Vertical (downwurd) 42
Lateral Type I aircrlft 25
Type II aircraft 30
Longitudinal (forward)
cockpit so
Peasenger compartment 50
Rlesultant vector 50
*The downward, sldewrd., and forvward velocity componentsof the revultant velocity vector do not exceed 42, 30. and50 ft/lec repectively.
The crash test sequence began when the instrumentation recording equipment wasstarted, approximately 30 seconds prior to releasing the test helicopter. The releaseof the helicopter was initiated in the control room by a push switch that closed relaysand sent signals from the pyrotechnic power supply to the guillotine cable cutters inthe pullback harness. A second and third relay initiated external and internal cameracoverage. A lanyard system for the internal flight programmer initiated the lightsequence after approximate;y 1 second of free-fall time. When the helicopter haddropped to where the landing gear was approximately 1 foot above the impact surface,the lanyard was pulled to fire the swing cable harness pyrotechnics, which in turnseparated all harness cables from the test specimen and permitted free flight to timeof impact.
IMPACT
TO4 took place on 4 August 1976 with release at 1430:0.0381. Impact occurred at1430:1.9243. After initial impact, the test helicopter slid approximately 10 feet onthe facility's concrete apron.
Actual impact parameters differed from the planned conditions, as shown in Table 2.Nevertheless, the resultant impact velocity is representative of the 95th percentilepotentially survivable crash pulse.
Actual impact parameters were determined by analysis of the still sequence and motionpictures and by the Doppler radar data. Analysis of the pictures permitted determina-tion of the aircraft's pitch attitude at impact, its flight path, and its vertical velocity.The horizontal velocity of the helicopter at impact was obtained using data from aDoppler radar unit that wa- located on the impact surface approximately 200 feet infront of the impact point. Radar data was continuously recorded.
The drop sequence, as obtained from the 70mm ground cameras, is shown in Figures8 and 9 for the right-side and left-side views of the aircraft respectively. Figures 10and 11 illustrate the crash impact severity.
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Figure 10. Post-crash view right front.
Figure 11. Post-crash view - lolt rear.
27
DATA ACQUISITION
The impact severity and aircraft breakup resulted in 19 channels of data being lostduring the impact. Of these, twelve were lost due to cables being severed by fuselageand component breakup, five were lost due to separation from their mounting pads,and the two fuel cell pressure sensors were lost when the forward landing gear im-pacted the fuel cell end plate.
All interior cameras functioned as planned; however, after the crash It was discoveredthat three camera lenses were damaged. This damage was caused by the extremeshocks sustained by those cameras located over the landing gear structure. The impactdrove -he landing gear upward into the aircraft and into the camera mounts. Someof the flash lamps were separated from their base sockets upon impact, but the mag-nesium filament trontinued to burn and provide adequate light.
DATA REDUCTION AND PRESENTATION
The test data were recorded on FM magnetic tape and processed by NASAILRC.After the analog-to-digital conversion using a sample rate of 4000 samples per wcond,the data were filtered using a 100-Hz low-pass digital filter. The 100-Hz low-passfilter Is si.nilar to the class 60 filter described in the Society of Automotive EngineersRecommended Practice SAE J2119 for full-scale vehicle impact test data reduction. Allacceirrmeter data were numerically integrated twice, using the trapezoidal rule integra-tion method with a 0.00025-second time step. The resultant time histories were thenplotted to provide the following information:
1. Unfiltered accelerometer, deflection tube, strain gage, restraint displacement,pressure transducer, Doppler radar, and load cell data.
2. 100-Hz filtered accelerometer, load cell, and deflection tube data.
3. Velocities and displacements obtained from integration of deceleration data.
Zero time for all plots was 1430:1.800, or 1.7619 seconds after the time of aircraftrelease; the data were plotted for a duration of 0.600 second.
In the following structural and restraint experiment discussions, only a portion of thedata that was acquired from T-40 is presented. Plots of data acquired from all 126channels are available for review at the Applied Technology Laboratory, U. S. ArmyResearch and Technology Laboratories (AVRADCOM), Fort Eustis, Virginia.
28
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EXPE RIMENTS
STRUCTURAL
Purpose
The purpose of the structural crashworthiness experiment was to obtain data for usein the development of analytical tech iques that can be used to predict helicopterstructural behavior over a range oi crash impact conditions, and for use in the estab-lishment of design criteria for crashworthy structures that will minimize crash forcestransmitted to occupied areas and thereby reduce occupant injuries. Also, the effectof structural breakup on fuel cell integrity was sought.
Background
Analysis methods, such as KRASH, for improving crashworthiness of helicopterstructures are in their infancy, but the correlation of these methods with testing isbeginning to lend credence to their effectiveness. Previous efforts to develop theKRASH model have been directed primarily to the utility helicopti-. Analysis of astructure as complex as that of the CH-47 cargo helicopter presents a more difficultchallenge. Accordingly, the Boeing Vertol Company was awarded Contract DAAJ02-76-C-001, (Reference 1). The objectives of that contract were: (1) V simulatethe dynamic response of a CH-47A helicopter for a crash impact as defined in Table2, using computer program KRASH; (2) to currelate the predictions of KRASH withthe data obtained from this test (TA0); and (3) to formulate and recommend im-provements to KRASH. To ensure adequate correlation of the test results withpredictions, Boeing Vertol specified the types and the locations of the structurallyrelated instrumentation that was to be installed in the test helicopter.
Description
To provide adequate data for correlation with the structural behavior predictionsobtained from KRASH, 34 strain gages were discretely mounted in the test helicopter.Also, floor- and component-mounted accelerometers and deflection tubes were installedfor correlation purposes. The deflection tubes provide a time history of the verticaldisplacement of the fuselage crown with respect to the cabin floor at six locations inthe cabin. Two other deflection tubes were rigged for post-test measurement only.
The test aircraft was equipped with two new standard (noncrashworthy) fuel cells.All connectors had been removed and connecting points securely sealed. Each cellcontained 1905.5 pounds of water. This weight is equivelent to that of 300 gallonsof JP-4 fuel, which is the design capacity for the CH-47A fuel cell. To measure the
BadriNath, Math Model (KRASH) - CH-d7A Crashworthiness, Boeing Vertol Co.,i •USARTL-TR-78-24, Applied Technology Laboratory, US Army Research and
Technology Laboratories (AVRADCOM), Fort Eustis, Virginia, to be publi.hed.
II 29
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hydraulic ram at impact, a pressure transducer was located at the forward bottom con-nector mount of each fuel cell end plate. The water was dyed red and green in theleft and right side cells respectively.
Results
A detailed discussion of the structural experiment results is contained in Reference 1.For this reason, only a limited presentation of results is provided in this report.
The nose-down crash attitude resulted in the forward landing gear making initial groundcontact. As the impact progressed, the gear rotated aft, impacting and displacing theend plates of both fuel cells. This reaction, coupled with pod and fuse!age materialfailures and resultant jagged metal punctures, resulted in a massive rupture of the fuelcel's and immediate dissipation of their contents. This effect is depicted in Figures 12and 13.
Figure 12. Fuel cell rupture and water dissipation.
30
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Figure 13. Post-crash view of fuel pod and cell.
The pres:ure transducers were destroyed by the impact of the landing gear on the endplate, but pulses were recorded for approximately 0.080 second after touchdown. Thesedata show hydraulic ram peaks ot 200 psi for the right -side cell and 170 psi for theleft-side cell.
The very rigid attachment of the landing gear to the box beam structure at station 240caused this structure to follow the gear rotation, breaking the aircraft at this location.This sequence is shown in Figures 8, 9, and 14. The interior floor deformation resultinqfrom this structural failure was severe, as shown in Figure 15.
*1 The transmission and engine mounts adequately retained these components, preventing
their penetration into the helicopter's occupiable interior. As an example, a failed realengine mount is shown in Figure 16.
31
S; ".-: - / . .... '- " ' . 'A.' " .-- -- -, -
- - I
Figuro 14. Structural failure at station 240.
•I IFi•3ure '3, Inteior tloo deformaioatttin4.
-39
I. IrI II I
I |r
-
Figure 16. Rear engine mount failure.
Standard CH-47A seats were installed in the test specimen. As shown in Figure 17,both seat pans broke as a result of the crash impact. While the failure of the copilotinertia reel allowed the dummy to pitch forward, the pilot inertia reel engaged, causingthe restraint to tear the back of the seat. This is also shown in Figure 17.
Figure 17. Seat pan failure.
33
Figure 18 shoyw the variation of the 100-Hz fift.red peak decelerations (measuredalong the cabin floor centerline) and the cabin height reduction with the aircraftstation.
3113
U 0
izlIIs$ 2ll 3ll 419 too
i' lI$|IRAi STATION (l1|
'r Figure 18. Variation In peak decelerations and cabin height
ireduction with aircraft station. J34
UIIL&6 SIAIIWI(II
In Figures 19 through 28, the ordinate spe.cifies the tape recorder/recorded channel andthe unit of measurement. (For example: A02-G indicates this is a plot of measuredacceleration (G) recorded on Channel "2" of recorder "A".) The abscissa depicts thetime history beginning at 1430:1.800. The Impact sequence began 0.1243 second later,when the forward gear contacted with the facility concrete apron. Each data plot hasa label that describes the instrumentation and provides other information as depicted bythe following example:
a b c d e ICH-7 Engine-Right LO R-496 5 N
a. Designated data channel number (1 to 126)
b. Transducer location, description, or purpose
c. Instrumentation type and/or function:
V - Accelerometer, vertical directionLO - Accelerometer, longitudinal directionLA - Accelerometer, lateral directionDTP - Deflection tubeLC - Loal cellSG - Strain gageSGI Strain gage, innerSGO - Strain gage, outerPT - Pressure transducerPTN - Navy pressure transducerSGN - Navy extensiometer
d. Aircraft location Icenter (C), left (L), or right (R)l and fuselage station
e. Transducer sensitivity X 10-2 G
f. Recorder: Army (A), NASA (N)
The deceleration time history during impact for the engines and transmissions is presentedin Figures 19 and 20 respectively. Comparison of these data with Figure 21, which givesthe cabin floor and landing gear attachment deceleration time history, illustrates the energy-absorption capability of the deforming structure. For instance, deceleration readings in thevertical direction range from 210.5 G at the rear right landing gear at station 482, to59.8 G at the floor centerline at station 460, to 9.2 G at the rear transmission atstation 540.
Figures 21, 22, 23, and 24 deal with the deceleration time history of the cabin floor andreflect the extent of the data reduction accomplished. This includes:
* Raw and 100-Hz filtered data.
* Expanded scale filtered data.
b 35
* First integration of data for velocity.
* Second integration of datd fur displacement.
Figure 25 gives the time history recording of cabin height reduction measured by thedeflection tubes. Figure 26 pr-sents the longitudinal deceleration time history at thecabin floor centerline for various stations. Also shown is the fuel cell hydraulic rampressure at impact. These pressure transducers were destroyed just after impact.
Crash force attenuation from cockpit flocr to seat to occupant is shown in Figures 27and 28 for the pilot and copilot stations respectively.
CONCLUSIONS
1. The data acquired from TO4 provided a sound basis for empirical correlationwith the computer model KRASH predictions and subsequent improvement of
the model.
2. The rigid forward landing gear attachment caused aircraft breakup at station240 in both T-39 and T-40. This structural deficiency poses a crash safetyhazard to cabin occupants, which could be eliminated by designing the attach-ment to fail at a load lesser than that necessary for structural member failure.
3. Seat pan failures at both the pilot and thp copilot stations are indicative ofthe severe tailward decelerations experienced by occupants in a 95th percentilepotentially survivable impact. :Had this been a manned helicopter, it is mostprobable that the pilot and copilot would have sustained incapacitatinginjuries.
4. The interaction of the forward landing gear with the fuel cell end plate forthe nose-down crash attitude poses a potential CH-47 post-crash fire hazardwith or without crashworthy fuel systems installed.
3
b 3
IT
I H 444- 4 1-+12j- H-1 -H44Ifflifff M tvt tj I
CH I XMSN41D LO C-110 5 N 7
t ffýff,. . ............ -----
-1 URFP 14
H 1 4-TtEth#1 ý4' CH 2 XMSN-FWD V C-110 2 N -ii
t ...... t il 4ýt
r t qTT 4
T T- I- tl
H+H1141f# _ 11+- ----------
CH 3 XMSN-REAR LO C-540 5 N #.............
.... .........L1 L
Li
AFa H-%"
+
CH 4 XMSN-REAR V C-540 2 N . .....
........ --- ------........... Im;
7 -------- +
CH 5 MIXER-BOX LO C-455 1+ 114.17
7-
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L] U I7f _T L T
f 44+
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Figure 19. Gearbox decelerations.
37
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-71
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t+ 4T t f,+f _i_ý4_4+1 f4+4 i I
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4- 141+ _L1.1A
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Figure 20. Engine decelerations.
39
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~t .j ~ H 5 FLOOR V C-165 7 N -
I IV
CH 21 FLOOR V C-370 10ON _____
..... . . . . . . . . ..... .. ..! . ..
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.; C H 19 FLOrR V C-460 7 N
i"t 4 l IIL .L 211t-
44T CH 13 LANDING GEAR-R REAR V R42 .ON .... ____
.,4 U.CH 14 LANDING GEAR-LREAR V L-482 10 N K
I'CH 15 LANDING GEAR-R FRONT V R-240 10 N ....
If I
L_ II~ ix~..r~j~h.:.............
Figure 22. Cabin floor and landing gear attachment vertical decelerations(100 Hz filtered).
I 43
tII Im p h !!I fil fi l I l l 1 1 f l i l Ill ýf l i f 1 1 1 1 1
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CH 25 FLOOR V C-250 7 N i
CH 21 FLOOR V C-370 I0 NT
SCH 19 FLOOR V C-460 7 N .It
St-tl it tif t!t ii ' 1 i ~ ~ t! :• [, ; i' 1;j
CH 13 LANDING GEAR-L REAR V R-482 10 N i
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p T7 I pl,! 1, " f.ý 77,+t t t71+1 1 Tilt I
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R 240 10 N i t::;CH 15 LANDING GEAR-R FRONT V .. ... ... ...
I .. .... ..
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T + fl- I 1 1
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Figure 24. Cabin floor and landing gear attachment vertical decelerations
(diL.placement integration).
47
N
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____ ___ ____ I ! i~ d t ik i I ___w 11 1
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4 CH 61 DEFLECTION~ TUBE OPT L-455 N
Tll t
CH--64 DEFLECTION TUBE OPT L-245 N
CII 65- DELCTO TUB DP i. 2 N M
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T__ CHI 17 FULO CELLO LEFT PRSSR 7-240
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tit 1+1111iiiiiIH I -i H 141J: CH 48 SEAT FLOOR-RIGHT LA R-075 1 N
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+
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H
-RIGHT LO R-080 5 N 1 4 44
CH 52 DUMMY r
J..11:H Itf i t ........... ...
CH 53 DUMMY-RIGHT LA R-080 2 N N4 tý +
-t A
N
... ... -- --- --- '-T ýt i 14f
ýit Aih -t* lif
CH 54 DUMMY-RIGHT V R-080 5 N
M iffý -4V Wt1ý11 1 '111 It I't It 141 tit 41H t1t Mt. H 11 tt iý 11 tf 41 i4 14 p p p p F p I I Iii iiii
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tit 121 t-, t
Figure 27. Pilot station decelerations.
53
JV
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.4~0 VI t ., ~ *, 1~r.4
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Ci50 SEAT FLOOR-LEFT LO L-075 7 N~t { jj j
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Figure 28.T CopLOO-ET sttV deleiatIons
fittlI i l 1H5Oil
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44
CARGO RESTRAINT
The load attenuator or energy-absorber concept looks promisir-g as a means by whichinternal helicopter cargo loads may be rentrained with a minimum of individual tie-downdevices. By reducing the number of required tie-down devices over that normally required,the downtine necessary for a helicopter to be loaded or unloaded inder combat conditionsmay be significantly reduced, thus reduzing the exposure tima 3f the helicopter and crewto enemy ground fire. The load attenuators, fabricated especially for this test, were demon-strated under actual crash conditions for the first time during this test.
Purpose
The purpose of the cargo restraint expe-iment was to demonstrate the effectiver,e-s andthe performance of the wire-platten-type load attenuating device when it is installed in serieswith conventionrl cargo restraints in an actual crash environment.
Description
Five individual experiments were conducted aboard the CH-47 test helicopter: two M-100'h-ton trailers with simulated cargo, ed(L,, with a nominal weight of 1,315 pounds; two palletloads of sirmulated 155mm howitzer projectiles weighing approximately 800 pounds each;and a cargo package ;,r baseline crash data acquisition, which was rigidly mounted to thefloor at the ai,'craft longitudinal CG.
All items of cargo were positioned and restrained according to TM 55-450-18, CH-47 Inter-nal and External Loads Manual, by experienced personnel from the Fort Eustis Transporta-tion School using the standard MB-1, 10,000-pound-capacity chains and the CGU-1/B 5,000-pound-capacity nylon straps. The final configuration of the cargo and restraints is shownin Figure 6.
The bpseiine crash data package consisted of a free mass attached to two load cells thatsensed forces in the lateral and longitudinal planes. In addi on, accelerations in all threeplanes were measured by accelerometers. The package that was located on the stdrboardside of the carqo compartment is shown in Figure 6.
In the forward section of the cargo compartment, one trailer was restrained in the conventioncl manner with six MB-1, 10,000-pound-capacity chain restraint devices, and one palletload was restrained with four CGIJ 5,000-pound-capacity nylon strap devices. The remairing trailer and pallet loads were restrained in a similar manner in the aft section of thecargo compartment with the load attenuators connected in series with the normal restraints(Figure 29).
The load attenuators used in this test were ,he wire inding/extrusion type (Figure 30),whose dynamic characteristics can be controlled du: ,g the manufacturing process byvarying the diameter and metallurgical properties of the wire and the distance betweenthe extrusion rollers. The energy-absorbing devices that were ujsed in seriE- with the cargorestraints were designed with an initial exten-ion force (breakout force) of 4,940 pounds,which is just under the maximum strength of the aircraft floor fittings.
57
Figure 29. Rear pallet load looking forward, after Impact.
#eua~##title
flitpfl
l~ Twitil U. TJFigure 30. Cargo load attetnietor.
Load cells were placed in series with the forward restraints on all cargo. Triaxial accol-erometers were mounted on each load and on the floor beneath each load. Rapid sequencecameras were mounted at various points In the cargo compartment and focused on loads,restraints, and floor fittings to record the impact dynamics.
Results
After the CH-47 came to rest, all items of cargo were generally in their original positions,although some of the restraints had become detached by the end of the impact sequence.There was also some evidence that thle forward trailer impacted the bulkhead at the headof the cargo compartment (Figure 36).
Of the four items of cargo in the helicopter, the aft trailer equipped with load attenuatorsin series witn the restraints was the only item to sustain major damage.
The aft trailer experienced high vertical G loads as evidenced by the 130 G peak meas-ured on the floor beneath the trailer. These high vertical forces compressed the trailersprings, causing the shackles to penetrate the floor and to severely distort the axle, thewheel rims, and the chassis (Figures 31 and 32). The trailer rebounded from this positionand induced loads of sufficient magnitude simultaneously in all restraints to stroke theattenuators on both the forward and lateral restraints and to fail tlhe two aft restraintfloor fittings. The reason that the attenuators in series with the aft restraints did notstroke and prevent failure of the floor fittings is not clearly understood. The nonstandardattachment of the starboard floor fitting to the base of the crash data package may havecontributed to its failure.
The port floor fitting (ST-340) was subjected to tensile loads severe enough to cause itsmounting bolts to tear aftward through the floor for a distance of 7f8 Inch withoutstroking the load atter uator (Figures 33 and 34). In this case the integrity of the loadattenuator was questioned. Unfortunately, load cells were not available for direct loadmonitoring of the aft restraints. The forward trailer also exoerlenced a bent axle fromtvertical G loads of 80 G peak measured on the floor beneath the trailer, although itwas still serviceable after impact, as was most of the cargo. However, the crushed anddistorted fuselage of the helicopter prevented the immediate removal of the cargo withoutspecial equipment.
The right-hand forward trailer restraint, torminating at ST-240, sustained a nearly constanttensile load of 5,255 pounds for a period of .032 second before failure of the floor fittingdue to buckling of the floor (Figure 35).
?" 5 9
. .L-..
". '
Figure 31. Rear trailer looking forward from starboard side after impact,
Figure 32. Rear trai~er looking forward, after impact.
60
Figure 33. Rear trailer, aft restraints with energy absorbers.before impact (station 340).
"7"Z
Figure 34. Rear trailer, aft restraints, after impact (station 340).
61
j1,
4A
-7..- � '
I.7.,V ,1
-�- �
� 4 f.� 2 . N .- �d'�' 's;
� ,A:�.
'V* 7)a � � "�K; K
/
�1 .g______ .r *�
Figure 35. Forward trailer, forward and lateral restraints,after impact (station 240).
Figure 36. Forward trailer just aft of bulkhead.
62
pI
I
Post-Crash Laboratory Test
A laboratory test was conducted -it ATL in an attempt to learn why the floor fittings atST-340 failed without extension cf the load attenuators. The objective of the laboratorytest was to determine the exact load at which the two unextended atteruators, whichwere attached to the floor fittings at ST-340, would break out and begin to stroke. Withthis information, it could be assumed that the tie-down fittings were subjected to loadsjust short of this amount. A third unused load attenuator was stroked completely in 141milliseconds to simulate a crash pulse.
Both of the questionable attenuators were stroked at a rate of 4 inchec per minute,duplicating the manufacturer's performance qualification tests.
Under test conditions, both attenuators achieved peak loads of 5,200 and 5,300 pounds(breakout force) and stroked at level loads of 4,800 and 5,000 pounds, which was abovetheir design performance of 4,940 pounds peak load and 4,600 pounds stroking load(Figure 37).
The results of these tests suggest that both floor fittings at ST-340 could have been sub-jected to loads up to approximately 4,600 pounds, since some indication of attenuatorplatten wire deformation would have been evident at loads above this level. CH-47helicopter floor fittings have failed under static test loads of 5,050 to 5,150 pounds inthe past; however, the age, corrosion, and general deterioration of this particular CH-47fuselage probably contributed to early failure of these fittings.
The third test produced surprisingly similar results with a breakout force of 4,818 pounds,only 122 pounds less than the design value, which reinforces the validity of the manu-facturer's low rate method of testing.
Conclusions
In T-40, loads were developed in the cargo restraints which exceecied the ultimate strengthof some of the floor tie-down fittings in the test helicopter. Energy-absorbing devicbswere effective in attenuating peak loads in cargo restraints during the helicopter crash, asevidenced by the extension of both the forward and the lateral load attenuators on theaft trailer.
Noting the two failed floor fittings (ST-340), the results of this test suggest that thedesigned attenuator breakout force should be substantially lower to allow more energydissipation and earlier stroking to avoid any chance of excessive peak load buildup. Theload attenuators used in this test were designed with a breakout force of 4,940 pounds,which now appears to be too high for maximum system energy dissipation in this case,and perhaps too close to the ultimate strength of the floor fittings in the CH-47 helicopter.
In the final design of a load attenuator of this type, accuracies greater than ±400 poundswould be difficult •nd expensive to achieve, considering practical manufacturing tolerancesand the variances of metallurgical cl-racteristics of available materials.
63
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In selecting attenuators for cargo restraint, a much greater margin should be allowedbetween the ultimate strength of critical components and the limit load value than wasallowed in this cargo restraint experiment.
The load attenuators used in this experiment were designed to fit in series with the MB-1chain, or the CGU strap restraint devices; however, several operational problems wereexperienced that will require future hardware modifications. One of the hooks on theCGU strap restraints did not fit the atten,.uator body and a shackle was required. Thelong sections of wire extending out frum the attenuator body interfered with operationof the MB-1 latch and t!Ve tonsioning device and should be repositioned in the finaloperational hardware (Figure 38).
The results of this experiment have provided some valuable design data for future cargorestraint systems, although the test impact conditions were less than ideal for an internalcargo restraint experiment. Impact conditions having a higher longitudinal velocity com-ponent would produce a more challenging environment for evaluating cargo restraintcomponents.
The use of load attenuators in series with conventional cargo restraints improves thesurvivability aspects of a heliconter crash and reduces the number of individual restraintsrequired. As more emphasis is being placed on transporting cargo internally, maximiza-tion of the total load on each flight is essential to increase productivity of the helicopter.
The plan view of the test CH-47 cargo compartment (Figure 6) illustrates how the place-ment of conventional cargo restraints, radiating outward from each load to the floortie-down fittings to achieve the most advantageous restraint angle, limits the cubic capacityof the cargo compartment.
Development of an energy-absorbing system to facilitate quick restraint of large quantitiesof internal cargo in one operation would promote optimum utilization of the total cargospace.
Figure 38. Partially extended energy attenuator showing interference of platten wirewith tension adjustment wheel on MB-1 chain restraint.
65
'.
AIRCREW INFLATABLE RESTRAINT
Purpose
The purposes of this experiment were: to demonstrate that an aircrew inflatable restraintsystem is capable of functioning under "real world" conditions; to provide comparativedata between an inflatable and a conventional restraint under identical conditions; and toobtain the "crash signature," i.e., the magnitude and duration of the crash pulse, to aidin further development of the crash sensor.
Background
The Naval Air Development Center is developing an inflatable restraint system for helicop-ter crewmen, which provides greater crash proiection to the wearer than the standard lapbelt/shoulder harness system now used. The inflatable restraint provides automatic pre-tensioning, which removes any slack in the system and forces the occupant back in hisseat, thereby reducing dynamic overshoot. Strap loading on the wearer is reduced whenthe inflated restraint is compressed during crash loading. The concentration of the straploads on the body is reduced because of the increased bearing surface provided when therestraint is inf'ated. In addition, an inflated appendage of the restraint that fits underthe wearer's chin will reduce the head motion and the whiplash induced trauma whichoccurs during a crash. Submarining by the seat occupant is prevented by the use of acrotch strap.
The inflatable restraint system was designed using the air bag concept of enveloping theseated occupant with a gas-filled inflatable device to prevent fatalities and to reduceoccupant injuries during a potentially survivable crash. Unlike the automotive air bag,which is a passive device remotely located from the occupant, the inflatable restraintis worn in a fashion similar to the present-day crewman harness. The inflatable restrainthas been configured essentially the same as the harness currently being used in helicop-ters and other nonejection- seat aircraft. These harnesses usually consist of 1-3/4-inch-wide shoulder straps and 3-inch-wide lap belt straps joined at a central fitting. Bothshoulder straps are joined directly behind the occupant's neck and terminate in aninertia reel mounted onto the seat back. The ends of the lap belt are anchored to thelower rear portion of the seat bucket.
The restraint system (Figure 39) is comprised of three rn, ir subsystems: (1) the inflat-able bladder/restraint, (2) the inflator, and (3) the crash sensor. The system has beendesigned so that in its stowed position it appears somewhat like the conventional harness.When unfurled, the bladder/restraint is revealed as shown in Figure 40.
The inflator is a cylindrical pyrotechnic gas-generated inflator manuafactured by the ThiokolChemical Corporation for this system (Figure 41). It is located within the bladder.
66
INFLATEDUNINFLATED
TEAMIN 4FO INTO* ~AN INERTIA R~EEL
RELEASE DEVICEAND ADJUSTER
Figure 39. The inflatable restraint system.
Figure 40. Comparison of inflatable restraint in stoweL. position and unfurled.
67
/ / P
"Figure 41. Thiokol Chemical Corporation's pyrotechnic inflator.
A nontoxic gas, mainly composed of ,iitrogen, fills each of the restraint bladders to apressure of 3 psi in less than 20 msec after initiation. The inflator ignition system isinitiated by an electrical 7.ulse from a crash sensor located on the floor of the aicraft.The crash sensor is an acceleration switch selected because of its ability to act as anintegrator of acceleration time. The sensor is set to close at an energy level of 5 Gfor 11 msec. Inflation of the restraint occurs within the initial 30 msec after the sensorcloses. As the occupant's body moves into the bladders, the bladders' pressure increaseswhile they are being squeezed by the torso against the outer straps, offering furtherresistance to the occupant's movement into the restraint. The material, being semipermeable,then allows the gas to escape so that in less than 0.1 second after initiation the pressurehas decreased to its original 3 psi, and the bladder continues to deflate with time. Shoulda secondary impact occujr, the uninflated restraint is still positioned around the occupant,offering protection against further decelerative forces.
Description
To obtain the comparative data, two 95th percentile dummies were restrained in the pilotand copilot seats by an inflatable system and a conventional restraint system, respectively.Each dummy was clothed in white thermal underwear and wore a Navai aviator'q helmet.No slack was allowed in either restraint. Standard CH-47 seats and cushions w.,- used,and no attempt was made to improve the strength of the sext or its support structure.
Each dummy had a triaxial accelerometer mounted in the chest cavity, and each restrainthad force transducers mounted on the webbing to measure loads on the lap belt andshoulder harness. In addition, a pressure transducer was used to measure the internal gaspressure of the inflatable system. The cockpit was photographically covered by two high-speed motion picture cameras equipped with wide angle lenses. The crash sensor wasmounted on the floor in the rear cockpit area between the seats. An initiation wire ranfrom the gas generator located inside the bladder of the restraint through the crash sensorto a battery pack in the cockpit. A triaxial accelerometer was mounted near the crashsensor to record the acceleration levels of the crash.
68
/
Results
A review of the motion pictures revealed that the inflatable restraint system functionedproperly during the crash. Although the seats were considerably deformed by the crashforces, the inflatable restraint was able to constrain the dummy in the pilot seat (Figures42 and 43). Unfortunately, the inertia reel on the copilot's seat failed, releasing theshoulder straps end allowing the dummy to pitch forward in the seat. This unforeseenfailure eliminated all chance for a meaningful comparison of the data collected fromtransducers on both dummies.
Drop test data are presented in Table 4. Time response plots of the data summarized inthis table are shown in Figurws 44 through 52. All data shown ;n these plots werefiltered at 100 Hz.
Analysis of the data showed that the pressure developed in the inflatable restraintreached a maximum of 9.G psi. This is considerably iess than was experienced duringhorizontal sled tests, but it is explaincble since peak pressure results from the compres-sion of the bladder as the occupant's torso loads the stra:js. For T-40 impact conditions,the motion is directed predominately downward into the seat bucket. The severedeformation of the seat absorbed a pordion of the energy, resulting in reduced strap loadsand lower internal pressure. Figure 43 shows the failure of the seat pan and the tearingof the seat back due to the strap loading; it aiso reveals the inflatable restraint in apartially inflated condition because of its semrporous nature. This feature permits aquick removal of the restraint by the wearer and a rapid egress from the aircraft.
S'* '* 71> -
Figure 42. CH-47A test vehicle (post-test).
69
INN
Figure 43. Post-test cockpit showing crash damage and dummy position.
TABLE 4. SUMMARY OF DROP TEST DATA
PeakLoadTimeParamneter Psi G Lb (mnsec)
Interne Pressure 9.6 106Floor, Gx 19.0 - - - -68
Floor. Gz 158.0 58Pilot Chest, Gx 12.6 81Pilot Chest, Gy 3.7 62Pilot Chest. Gz 15.9 55Pilot Chest Resultant. Gr 19.2 106Copilot Chest. Gx 9.6 94Copilot Chest. Gy 6.0 105Copilot Chest, Gz 21.7 61Copilot Chest Resultant. Gr 22.7 61Pilot Left Shoulder 535 illPilot Right Shoulder 666 103Pilot Left Hip 385 115Pilot Right Hip 247 176Copilot Left Shoulder 158 156Copilot Right Shoulder 135 62Copilot Left Hip 379 188Copilot Right Hip 394 ISO
70
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71
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72
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73
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74
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Figure 48. Pilot and copilot resultant chest acceleration verwus time.
75
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Figure 49. Pilot horizontal chest acceleration and bladder pressure versus time.
76
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Conclusions
TAO0 demonstrated that the aircrew inflatable restraint system can function under actualcrash conditions.
The crash pulse magnitude and duration obtained can be used in the development of acrash sensor for the advanced engineering model of the inflatable restraint system.
Though data permitting direct comparison with a conventional restraint was not obtained,the observed bladder inflation should result in lesser longitudinal forces being transmittedto an occupant equipped with an inflatable restraint.
80
APPENDIX AINSTRUMENT'ATION
GENERAL
T-40 measurements were obtained using five basic type transducers: strain gages, accel-erometers, load celis, displacement sensors, and pressure sensors. The locations of thetransducers on the test helicopter are depicted in Figure 6. All transducers were connec-ted to remote recording equipment via multipair and coaxial cabling. Nine recorderswere employed, three of which were supplied and operated by the Army. The Armyconditioning and recording equipment was positioned in a 40-foot instrumentationtrailer configured to meet this project's requirements. The NASA equivalent equipmentwas positioned in their permanently configured drop test site control room and wasoperated in accordance with their usual recording procedures.
NASA and Army instrumentation technicians worked closely together in the installationand wiring of the transducers, and in sensor calibration and data acquisition.
APPLIED TECHNOLOGY LABORATORY (ARMY) RECORD!NG SYSTEM
The Army data recording system is graphically depicted in Figure A-i. The Army recordeddata from structural strain gages, strain gage type accelerometers and load cells, and straingage pressure and restraint belt sensors. This was accomplished using designated taperecorders 0, P, and N. Recording circuits are described below.
Structural Strain Gage Recording Circuits
Thirty-four Micro-Measurements type EA-13-125AC-350/W constantan alloy 350 ohm straingages were used in conjunction with Vishay type MDL 311218-X bridge circuit completionmodules to form the transducer sense circuit (Data Channels 77 through 110). The out-put of each strain gage bridge was fed into an lED type CSO-340 Sub-Carrier Oscillator(SCO) and mixed in an lED type CMA 400A Mixer-Amplifier to form a 5 to 6 subcarrierfrequency multiplexed signal. All 34 strain gage circuits were fed into seven SCO signalconditioners (Modules G-M) with the resulting outputs fed via individual 200-foot 52 ohmcoaxial cables -o the instrumentation trailer where the signals were recorded at 60 ips ona 14-channel Gz3nisco megnetic tape recorder (Rec N), Model 10-126. A 100-kHz record-ing reference signal was also injected into each of the seven mixer-amplifiers to becomepart of the seven multiplexed signals. Recorder N also recorded IRIG A time code andsupplemental voice signals. Circuit calibrations of all ATL recorded channels were accom-plished on test day and consisted of circuit zeroing, plus a strain gage/bridge shunt typecalibration. The selected calibration shunt resistance value was equal to 16065 pu inch/inchtension. This was equivalent to approximately 50% of the FM recording band edge. TableA-1 presents a description of recorder N functions and calibration data.
Load Cell Circuits
Two BLH SR-4 10,000-pound load cells (Channels 75 and 76) were used in conjunctionwith the same carno package that was used in the March 1975 CH-47 (T-39) crash test.
81
*1.
This same package was used to establish comnarison load imoact data between T-39 andT-40. The load cell's 3ensor strain gage bridge output was recorded as part of the multi-plex arrangement described above. Calibration consisted of bridge circuit zeroing and aresistor shunt type calibration as determined by supplied data. The selected calibrationvalue was equal to 6,000 pounds tension. Eight iVCB Model 226A Quartz Force Linktype load cells (Channels 67-74) were placed in series with the rear restraint tie-downsfor all four on-board cargo loads: two 3/4-ton two-wheel vehicle trailers and two simula-ted pallet loads. Calibration consisted of circuit balancing and then inserting a knownvoltage equel to a known load cell output. See Tables A-1, A-2, and A-3 for the calibra-tion compression for a given channel. Refer to Figure A-1 for a general description ofthe load cell c&rcuity.
Pressure Sensors
Three pressure sense circuits were employed using CEC strain gage type Model 4-313and 4-326 sensors. Two 4-313 transducers (0-200 psi), Channels 116 and 117, wereinstalled in the lower front portion of the main fuel cells at Station 270. The thirdtransducer (0-25 psi), Channel 115, was employed to sense pressure exerted upon aninflatable bladder which was part of the dummy pilot's restraint system. This circuitwas a U. S. Naval Air Dev'.lopment Center (NADC) requirement and employed tech-niques and hardware developed by them.
Accelerometers
Two CEC Strain Gage Accelerometers, Type 4-202 (Channels 123-124), were used tosense center cockpit floor impact accelerations. The vertical sensor was a 0-250G com-ponent, while the longitudinal was 0-100G. These circuits, which were a NADCrequirement, were conditioned by the Accudata Model 118 and recorded on tapes 0and P.
Restraint Belt Sensors
'Eight NADC-fs.bricated strain gage sensors were employrd in conjunction with a stan-dard flight harness plus the inflatable restraint described above. Both pilot and copilotdummies had two hip and two shoulder tension sensors installed. The strain gage outputwas converted and calibrated to pounds of harness pressure using NADC methods. Foursensors were converted to an FM multiplex signal (Channels 111-114) and recorded onRecorder N while the remaining four (Channels 119-122) were recorded on Recorders 0and P, after Leing signal conditioned by an Accudata Model 118 amplifier arrangement.Calibrations were accomplished using selected shunt resistance values which producedsignals equivalent to known restraint belt pressures.
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NASA RECORDING SYSTEM
Figure A-2 generally depicts the NASA data recording system. Two different types oftransducers were recorded, piezoelectric accelerometers and extensiometers, on designatedrecorders A, B, C, D, E, and F.
Accelerometer Circuits
NASA recorded 62 channels of vertical, lateral, and longitudinal aircraft structural, anthro-pomorphic dummy and loads acceleration data using Kistler Model 818 and PC8 Model302A04 quartz type 0-1000G sensors with built-in impedance converters. The high-signal,low-impedance output (Nominal 10 mv/g) was fed through junction boxes directly into aNASA-fabricated unity gain impedance matching amplifier and ultimately into six Sangamoand Ampex 14-channel FM record magnetic tape recorders. IRIG A time code and testvoice narrative were fed into each recorder. Calibrations were accomplished prior to testday by injecting a kncwn AC RMS 1-kHz signal level at the accelerometer output (AircraftJ-Box) and setting the electronics gain as measured at and recorded by the respectivemagnetic tape recorder. Calibration levels were established at the 100, 200. 500, 700, and1OOG level depending on anticipated individual accelerometer G load. These calibrationpoints were established as 100% FM record band edge. Recorders A, B, C, D. E, and Ffunctions and calibration data are presented in Tables A-4, A-5, A-6, A-7, A-8, and A-9,respectively.
Extensiometers
Eight extensiometers (deflection/displacement sensors), six of which were electronicallysensed, were fabricated, installed, and calibrated by ATL personnel. The extensiometer isbasically an 8-foot 2-inch aluminum tube on which a nichrome rcsistance wire and wiperarm is side mounted over its length. The tube is secured to the cabin floor and thewiper arm to the fuselage exterior top. As the wire wiper traverses due to cabin floor oroverhead skin movement, an electrical current change proportional to displacement is pro-duced, allowing a time history to be recorded. In addition, a heavy rubber O-ring waspositioned on the tube at the roof line to sense the peak ceiling movement manually. To
. facilitate calibration and obtain optimum circuit sensitivity and linearity, signal conditioningcircuitry was fabricated. Calibrations were performed by creating a known deformationand measuring the resulting output voltage. Extensiometer positions are shown in Figure 6and physical particulars are depicted in Figure A-3.
Radar Velocity Measurement
A field-portable Doppler radar system was used as a signal source to determine aircrafthorizontal velocity at impact. Circuit calibration was accomplished using a tuning forkgenerated pulse equivalent of 20 mph. Data was recorded on Tape E.
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INTEF•'OR PHOTOGRAPHIC CONTROL
Th,- c',Y::-itry used to coo,-;ji Zhe test helicopicr interior cameras and liit s d-\sv-gn•dand fu,;'-.ated by ATL. This on~board carme.,a/iiht syctem is shown in Fi',.Ijre A-4.Camera 7-u 1*3ht locationc; are shown in Figure 7.
DATA AGOUMTbON
Procedures
All channels were caiibrated as shown in Tables A-1 throuch A-7. A i& ,z s:'.:Il wasrecorded on each tape recorder for sp.eed comnpensation and noise-reduction purpeses.Timing information waY_ recorded on ejch tape to facilitate data rteduction end t~vent occur-renc• definition. Also, an audio track was included on each tape recorder (edga or distincttrack) to provide for tape data marking, f,tcilitatimg more rapid data reduction.
Data Recording Problems
Some data was acquired on every channel. The following instrumentation malfunctionswere a direct result of the crash impact.
1. Data Channels 74, 82, 83, 89, 91, 102, 116, 117: Cables severed by struc~uraldamage.
2. Data Channel 92: Cable or transducer cut or smashed (sensur and completecable run could not be checked due to fuselaga coilapse).
3. Data Channels 15 and 18: Accelerometer separated from mounting pad dueto epoxy bond fracture.
4. Data Channels 13, 14, 49, ,nd 51: Accelerometer separa.-d from mounting
pad due to strippo.d pad threads.
5. Data Channeis 31, 32, 33 and 34: Cable severed due to structural darne.
AIRCRAFT INSTRUMENTATION MOUNTING AND CABLUNG
All aircraft test cabling was either RG-174/J Type 50 ohm coaxiid (FSN G145-jC3-3237)or Belt en Type 3434 four-corductor shielded (FSN 6145-011-2.20) and was looselypositior.,%d with extra slack at expected heavy stress points. All CC volt-:v and signalcircuits u-ed. the Type 8434, and the 1CO-kiiz plus the frCquecncy multip!z3xred strain gagesignals were distributed using the RG-174/J. Plastic type cable tie clcmps of variouslengths along with masking tape were used to suspend and sccur2 rast cM3 ling.
The seven SCO/Bridge module (G-M) circuitry mounts consý,icd of tv..o E,/'32- x 4-1/2- x11-inch aluminum plates separated by two 4-inch plastic sponr2 ssýpciaritors, oil of whichwere secured together through four holes in the fusel•a•e top using four 1 4-inch solf-locking aircraft hardware. One plate and one sponge section were rnounted on each sideof the fuselage skin.
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The seven lED SCO mounts (G-M) were aecured to their associated module mounts using10,32 self-tightening aircraft type hardwafe. The Vi',hay strain gage completion moduleswere adhered to the module mounts using Dow Corning Silastic 732 RTV silicone adhesive(FSN-8040-8658991). This adhesive was selected to allow undamaged module recoveryupon completion of the test. The module board terminal strips were bonded using 3MScotch-Weld Brand Epoxy Resin Structural Adhesive, Type 2216B/A (FSN 8040-145-0432).
All strain gage transducers were bonded to the aircraft structure using standard surfacepreparation, degreasing, plus neutralizing methods and practices as recommended by manu-facturers such as Micro Measurements Inc. (MM). The MM strain gages were fingerpressure adhesively bonded using Loctite Type 06 Super Bonder (Eastman 910 equivalent)and moisture protected with MM Type Gagekote 3 varnish. The sensor output wascoupled to the associated bridge completion module using a Belden 8434 cable conductor.
All acelerometer mount blocks were fastened to their mounting positicni. by 1/4-inrhor 3/8-inch hardware or adhesive bonding using 3M Scotch-Weld Epoxy Resin StructuralAdhesive, Ty.* 2216B/A. The accelerometer mounting stud bases were bonded to themount block using the same epoxy adhesive, and the accelerometers were secured to thebase using a 10-32 stud torqued to approximately 18 in.-Ib. The mounting techniques arethose specified by the respective accelerometer manufacturer. A Belden 8451 two-wireconductor or equivalent was used to connect the transducer to the respective J-Box.
Conventional mounting was utilized for each load cell in accordance with manufacturers'specifications and dimensions.
The camera/light control circuitry and distribution panel were prefabricated in the labusing a 1/4-inch aluminum plate as a foundation and securing all components with RTVsilicone adhesive or appropriate size hardware. AWG 10, 12, 14 and 18 gage wire wasused to interconnect control components and feed the light and camera loads dependingon current requirements. A pre-drop system check was performed to insure properoperation and time sequencing.
The ATL and NASA aircraft cable junction boxes (J-Boxes) were mounted at Station 330on the right and left Interior sides respectively. Both J-Boxes (with covers) were con-structed of 1/16-inch or 1/8-inch steel and were mounted to the aircraft skin using 3/8-Inch self locking aircraft hardware, L brackets, and exterior reinforcement washers. Re-quired rows of terminal strips were either crew mounted using 10-32 hardware or adheredusing Devcon Quick Set Epoxy Resin Adhesive, Type 5 minute. All attached cabling wasfirmly bonded for strength and secured at each box entrance. Connectors were con-ventional BNC or aeronautical MS circuit type. See Figures A-5 and A-6 for Army andNASA Junction Box positioning without covers. The aircraft-to-control room umbilicalharnesses were routed through a window-moitnted flexible rubber fixture. Figure A-7shows the interior and exterior cable securing techniques. No crash impact damage wassuffered by eithctr J-Dox, umbilical harness, or any of the seven module mounts or signalconditioning units mounted above the aircraft waterline.
98
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Figure A-5. Army junction box.
Figur,9 Al6. NASA junction box.
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