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Deathblade said:

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an Army Apache helicopter crashed during a training mission at Fort Rucker,
Alabama, resulting in fatal injuries to the rear seat pilot and survivable injuries to the front seat
copilot. Figure 1 shows the seating configuration of the two pilots. U.S. Army Safety Center
(USASC) investigators at Fort Rucker, Alabama, reported the aircraft damage assessment and
aircrew member injuries. U.S. Army Aeromedical Research Laboratory (USAARL) researchers at
Fort Rucker, Alabama, concurrently examined the helmets, restraint systems, and crashworthy seats.
Crash kinematics were derived from the investigation; including estimates of the motion of the
harnessed occupants during the crash.
Summary of the accident investigation
While hovering at approximately 200 feet above ground level, a flight system component
failed. Corrective actions by the pilot produced a near vertical flight path, impacting on relatively
hard, dry soil in an estimated 5-degree nose down attitude, an estimated 5 degrees of roll toward the
left side. The terrain sloped downward and forward about 5 degrees. The aircraft left ground scars
that indicated a displacement down the slope of 5 feet from the initial impact location, with a
forward displacement component of I foot. The aircraft landed nearly flat and remained upright on
the nearly horizontal surface. These measures indicated the descent was primarily vertical. We
determined the motion of the occupants during the crash could be simulated by a computer model
given the simplicity of the crash kinematics.
The AH-64 Apache airframe is designed with three energy-absorbing components (deformable
tires, collapsible landing gear, and crushable fuselage underbelly) that are intended to dissipate and
lessen impact forces before they reach the helicopter floor. Ground deformation and fuselage
fracture are two factors that contribute significantly to energy dissipation. From measurements of
the distortion in energy-dissipating components and from experimental data of helicopter crash
tests, investigators estimated the energy absorbed. The peak vertical, forward, and lateral
accelerations were 41 Gz, 5 Gx and 2 Gy at the floor below the front seat with a helicopter initial
vertical velocity at impact of 50 feet/second. On impact, the helicopter developed a near vertical
fracture between the front and rear seat compartments. The rear portion, which contains the bulk
of the helicopter mass, was lowered by about 8 inches relative to the front. The rear of the helicopter
had 8 more inches of travel than the front, reducing the vertical Gs transmitted to the floor of the
rear seat. Since the front seat pilot survived and the rear did not, despite less acceleration forces,
USAARL investigators searched for other less obvious causes of death. Our attention eventually
focused on the helmet, the glare shield, and the retractor of the shoulder straps of the harness.
Unfortunately, the fracture caused the glare shield of the rear seat instrument panel to be pushed up
8 inches from its normal position in the cabin, and into the strike envelope of the helmeted rear
seated pilot, as diagrammed in Figure 2.
3
The copilot had a burst fracture to L3 lumbar vertebrae. This fracture is a major injury, but
survivable. Because he sustained only minor head injuries despite the presence of the optical relay
tube sighting equipment very close to his head, the automatic retractor was assumed to have locked,
preventing any significant reeling of the shoulder belt.
The autopsy report of the rear seat pilot indicated death was a result of a basilar skull fracture.
Approximately 12 inches of harness belt were spooled out, calling into question whether the inertia
reel retractor, designed to lock at 3 G of belt acceleration, had functioned as intended. The pilot
sustained no spinal injuries, providing further evidence that he was subject to less acceleration forces
than the copilot.
An examination of the pilot's helmet revealed a set of scratches parallel to the helmet visor
release knob, and evidence of an impact to the helmet near the limit of the knob range of motion.
The knob was left at its uppermost position on the slide track (see Figure 3). This led to the
hypothesis that the pilot's helmet impacted and became wedged under the glare shield, as displayed
in Figure 2. Further forward motion of the neck and body, while the head essentially remained
motionless under the glare shield, transmitted a force system of magnitude, direction, and duration
large enough to produce the basilar skull fracture. In our study, we test this hypothesis by using
crash simulation software, enhancing our understanding of the kinematics of interaction between
the pilot and cockpit during a helicopter mishap.
The biodynamic simulation
Occupant biodynamics were reconstructed using the DYNAMAN program (GESAC, Inc.,
1992), a software package which is based on the articulated total body (ATB) computational
simulation software (Fleck, Butler, and Vogel, 1975; Obergefell et al. 1988). Given as input a
number of body segments connected by mathematical representations of common joints, the ATB
automatically formulates the ordinary differential equations that govern the body motion. For a
given set of initial positions and velocities of these segments, ATB integrated those equations and
provided time-histories of the kinematics which then are post-processed to generate graphical
representations of motion. The software produced time histories of plots of forces, displacements,
velocities, and accelerations of body segments which are used to predict injuries. Head and neck
injury criteria were used to predict pilot injuries, and the dynamic response index (DRI) applied to
predict spinal injuries to the copilot. The predictions were compared to artual injuries received by
the two occupants. Results of the ATB simulations were used to draw conclusions about occupant
motion and possible injury mechanisms and assess the safety equipment performance.
The initial impact velocity and the time history of the accelerations of the helicopter cabin
floor are necessary input to the software. These data are derived from the site investigation of the
damaged aircraft, followed by an analysis based on the work energy theorem to account for energy
4
losses during the crash and simple kinematic equations for rectilinear motion of a mass. The landing
can be likened to collapsing a contractible telescope or dropping a spring-mass-damper system with
the larger stiffness and damping rates closer to mass and further from the point of impact (Figure
4). At the onset of the impact, fires and landing gear collapsed, followed by the crush of the fuselage
underbelly against the ground, and finally a fuselage fracture. Each collapse mechanism removes
vehicle energy while reducing the Gs transferred to the helicopter floor. This simple model works
well for the rear seat occupant because the bulk of the helicopter mass is behind the fracture.
Experimental crash test data allow the approximation of the floor vertical acceleration as a
constant during tire and landing gear collapse and as a triangular profile during fuselage crush,
fracture, and soil depression (Coltman et al., 1989). Given the kinematic relations using the symbols
a, v, s, for acceleration, velocity, and displacement, respectively:
fdv = fadt and fds = fvdt
Assuming constant acceleration over time from initial conditions to= 0 sec and vo (as yet unknown
vertical impact velocity), and where A is the acceleration in G and v, is a velocity (ft/sec) and less
than va at t,, it can be shown (Zimmennann, and Merritt, 1989) that the end of the pulse occurs at
tj defined by:
V0 - V1
- 32.2A
Also, where sA is the distance of collapse of the tires and landing gear (46/12 = 3.83 feet).
S2- vo2 v-
A -4. 1(2)
64. 4sa
Similarly, integrating the kinematic relations for a triangular pulse between tj and t2 yields:
t2- t1-3 2.~ -t - 32.2Aa (3)
and
1(4
= 32.2AX (4)
where s, is taken here as 1.5 feet of soil depression, fuselage crush, and fracture, and A., is the
peak acceleration of the triangular pulse. Last, expermental data indicative of the energy that is
5
absorbed during landing (or equivalently, work performed to deform soil and helicopter and to fully
collapse the landing gear) is required to complete the ,stimation of the acceleration profile.
Assuming that landing gear alone can dissipate the kinetic energy of a 33 ft/sec vertical drop, then
the work energy theorem implies:
v0 -v2 = 332 ()
Accident investigators estimated the work or energy of the soil depression, fuselage crush, and
fracture as representative of a 35.57 ft/sec velocity change. Hence, since v2 (velocity at the end of
the acceleration pulse) is 0 ft/sec, then v, must be 35.57 ft/sec for the rear seat. From equation 5,
the initial velocity v0 equals 48.5 ft/sec. From equation 2, the acceleration plateau A equals 4.411
G. Equation 1 gives a time tj of 0.091 seconds. The peak of acceleration triangle, G. is 26.2
Gs from equation 4. Finally, from equation 3, end of the pulse occurs at t2 of 0.175 seconds. The
acceleration pulse time history for the back seat pilot is plotted in Figure 5.
Because of the fracture between front and rear, the triangular acceleration pulse at the front
seat floor had to be shorter in duration and more severe in magnitude. Since fracture occurred after
collapse of the landing gear, the acceleration plateau A, velocities vo and v1, and displacement s.
should be the same as the front seat during the plateau portion of the pulse. The displacement
during the trangular pulse was decreased to 10 inches, so s. is 0.833 ft. Substituting this value of
s. and v1= 35.57 ft /sec into equation 4, we obtain A.. = 47 G. From equation 3, the acceleration
pulse ends at t2 = 0.138 seconds, and the pulse is shown in Figure 6. Both Figures 5 and 6 have been
smoothed, and sloped at early time to account for a more realistic and gradual Gz acceleration onset
due to tire deflection. Forward Gx acceleration was taken as a similar profile to the Gz, but
proportionately reduced to reflect A. of only 5 G.
Once the acceleration pulse was transmitted to the helicopter floor, further attenuation of the
signal was provided by the seat cushion and the energy absorber between the seat and floor. The
energy absorber provided a near constant force characteristic versus displacement once a 3540-
pound force threshold was reached, as shown in Figure 7. This device would not activate until a
certain G level was transmitted to the seat and pilot. For the copilot, the seat stroked between 5 and
6 inches, whereas the pilot rear seat stroked approximately 1 inch. In both instances, the floor
upward displacement impeded the seats' downward motion.
The model included body segment properties which were determined by GEBOD, another
component of the simulation software (Gross, 1991.) The model also included a harness belt system
and retractor to lock the shoulder belts at a desired time in the simulation. In simulations of the rear
seat pilot, the glare shield was placed in its final position because it was anticipated that the glare
shield was in its final position relative to the occupant before helmet impact. A helmet of the correct
inertia and mass as measured in a laboratory was added to the head segment, which now is referred
to as the head/helmet segment.
6
Results of simulations
Two scenarios for the pilot's biodynamics were simulated to attempt to achieve agreement
with his helmet damage, glare shield deformation, and injury evidence. A third simulation of the
copilot also was performed. The three simulations are discussed as follows:
Simulation I: Pilot's helmet impacts the glare shield during forward motion
This simulated scenario, where the visor knob impacts the glare shield during the forward
motion of the pilot causing the helmet to be wedged under the shield, represents the conclusion of
the original accident report. Figure 8 shows the sequence of pilot motion, with the initial seated
position that corresponds to the onset of the acceleration pulse, i.e., at 0 milliseconds (ms). The
graphics in this figure include the knob on the helmet, and the glare shield in its final position. At
100 ms in Figure 8, the occupant leans forward with an increase of relative angle between head and
neck. In this scenario, the harness belts are locked at 140 ms in a effort to duplicate the approximate
position of the locked belt in the actual mishap. Just before 144 ms, the large acceleration spike
causes approximately 1 inch of deformation in the energy absorber. At 144 ms in Figure 8, the
head/helmet segment just is touching the glare shield for the first time. Contact between helmet and
glare shield continues as the glare shield rides up the helmet until the glare shield impacts the helmet
knob at 156 ms. Immedistely after impact, the knob rebounds off the glare shield in Figure 8 at 164
ms and head/helmet motion continues below the glare shield.
Between 164 and 180 ms (see Figure 8), the head goes through a rotation as the forces that
have developed in stretching the harness belts pull the torso back into the seat. Eventually, the
occupant would be pulled back by the harness belt to the upright seated position. Figure 9 is an
overlay of the head/helmet as it moves under the glare shield. As the head contacts the glare shield,
the angle between the neck and head increases, increasing the likelihood of serious injury. When
the knob impacts the glare shield, the angle between the neck and head increases substantially and,
most probably, the basilar skull fracture occurred immediately after the knob impacted the glare
shield. Other evidence which supports this scenario includes the absence of damages to the
instrument panel and helmet shell indicating an impact of some sort on the helmet at the knob.
The force components in the neck along the aircraft X and Z directions, shown in Figures 10
and 11, have a resultant neck force that exceeds 1000 pounds for the period from 160 to 165 ms,
a magnitude and duration sufficient to cause serious neck injury (Coltman et al., 1989.) The
simulation results are susceptible to stiffness and friction coefficient of the glare shield and helmet,
and the position of the glare shield itself These are all parameters which only can be estimated or
known within a relatively large range. Adjusting these parameters could produce simulations with
larger neck forces, but will still demonstrate the same qualitative response as this simulation.
Finally, the potential that inertia reels failed to lock at 3 G, as designed, was explored. A
recent USAARL study tested 110 inertia reels in the field at Fort Rucker and found 24.5 percent
failed to lock at the 3 G requirement (McEntire, 1992.) This suspected failure in the Apache mishap
was supported further by evidence of approximately 12 inches of nonspooled shoulder belt from
7
the pIWlots inertia red. Faiure 12a shows a time history plot of the Gx acceleration of the upper torso
relative to the seat bzA which is equivalent approximately to acceleration of the belt as it unspools.
The plot shows acceleration did not exceed the locking threshold until approximately 130 ms. The
locking of the belt was simulated to occur at 150 Ms. Note that at about 70 ms, the Gx may have
been as high as 4 G; however, the Gz at that time was about 12, indicating the torso was moving
down into the seat and tended to relax the belt despite the forward acceleration trend.
Although the ATB formulation used in these simulations was not designed to generate the
linear acceleration of the belt at it spools out of the inertia reel, the generated output was
manipulated to extract shoulder harness extension versus time. The output was numerically
differentiated to produce harness belt acceleration at I ms intervals. Next, the accelerations were
smoothed by a running average over 5 points (5-ms window), and plotted in Figure 12b. The
computed belt acceleration (Figure 12b) indicates the 3-G level was exceeded at 70 ms into the
simulation for about 10 ms. Beyond 120 ms, an acceleration pulse occurred and, most likely,
activated the inertia reel locking mechanism. The large acceleration at 150 ms occurred as the
harness locked normally in response to the crash forces. The plot also indicates the simulated belt
acceleration may have reached or exceeded the 3 G locking threshold at an earlier time (70 ms) than
had occurred in the actual accident. Keep in mind the inertia reel locking mechanism was designed
to trigger at the onset of a 3 G belt acceleration. Nevertheless, uncertainties in the estimated input
parameters make the results of the simulations just as uncertain. Thus, it appears there is no
conclusive demonstration that the seat belt should have locked any earlier than the simulated 150
ms into the onset of the crash. Several simulations were run in which the inertia reel was
perima locked. As expected, these runs demonstrated the torso was restrained from excessive
forward movement, thereby preventing injurious head/helmet contact with the glare shield.
Simulation 2: Pilot's helmet impacts glare shield on rebound
In this simulation, the only parameter adjusted was the position of the glare shield, which was
moved slightly from its position in scenario 1. Figure 13 shows the pilot's head/ helmet skimming
past the glare shield without significant force between glare shield and helmet. The knob impact
and helmet deformation on its forward pass by the glare shield were, therefore, relatively
unimportant, and so the knob was ignored here. The shoulder harness belt was locked at 140 ms
and, as the belt compressed the upper torso it pulled the pilot back and his head/helmet impacted the
glare shield during the rebound pass with a very large force and at the location of the impact damage
on the actual helmet. Because this simulation did not generate sufficient downward forces which
are generally associated with basilar skull fractures, this scenario likely did not occur.
Simulation 3: Copilot's biodynamics of motion
In the actual accident, the front seat copilot sustained a serious lumbar L3 burst fracture due
to the large vertical acceleration. The investigators concluded the harness belts immediately locked.
Since the seat energy absorber stroked about 6 inches in the accident, a very stiff spring that
8
activated for deflections beyond 5 inches was included in the simulation to limit the seat and energy
absorber motion. The results of the simulation given in Figure 14 primarily show head neck motion,
although relatively large forces were transmitted to the spine from the seat.
To correlate actual spinal injury of the copilot with the simulated, plots of simulated energy
absorber stroke and seat vertical acceleration (Gz) are given in Figures 15 and 16, respectively. The
seat stroked approximately 6 inches between 110 and 140 ms before the floor came up and limited
the seat travel. During this period, the seat acceleration in Figure 16 was limited to a tolerable level
of approximately 15 G. However, upon floor/seat impact, the acceleration level abruptly increased
beyond 40 G. The acceleration remained above 40 G for 8 ms between 144 and 152 ms. It is
accepted generally (Coltman et al., 1989; Coltman, 1986) that sustained accelerations beyond 20 G
are very lik ', to cause spinal injury. An accepted measure of the potential to cause spinal injury
is the dynamic response index (Brinldey, 1985) which was developed from experimental studies of
pilots in ejection seats, and is based on the following spring-mass equation model approximation
(Brinkley and Shaffer, 1971):
d 26 53 i
d + 25.3 -3 + 27986 = 386. 1A,, dt2 dt ,t (6)
where Ak is seat acceleration in in/sec. This equation must be integrated numerically because
of the tabular form of the right hand side. The DRI is calculated at each time step as:
DRI = 7. 2488 (7)
where 8. is the maximum value of 8 as obtained from integration of equation 6. This resulted in
a DRI equal to 23. This led to an estimate of probability of spinal injury rate of slightly greater than
50 percent (Brinidey, 1985). It might have been expected that this would have been larger because
of the seriousness of the copilot's injury. However, this result does indicate that injury would be
anticipated in this environment.
Several other simulations were performed where the seat stiff spring was removed, allowing
the energy absorber to fully stroke without impacting the floor. By allowing the energy absorber
to stroke a full 10 inches, it limits the acceleration to the tolerable 1 5-G design point. Severe injury
may not have occurred. It is reasonable to conclude that the interference between the seat stroking
and the usual floor motion was a significant factor in increasing the risk of spinal injury.
9
Biodynamic modeling confirmed the injury causing scenario proposed by an accident
investigation team, and refuted another suggested scenario. This type of simulation is a quick and
powerful tool that allows investigators to obtain reasonable estimates of the internal forces in the
neck and the lower spine which cannot be obtained by other means. The ability to simulate different
scenrios allowed us to conclude the stroking of front seat is all but lost in high impact accelerations
when the floor buckles upwards. Used with accident evidence, we can explore how safety devices,
e.g., inertia reels and energy,--.bsorbing seats, functioned in a crash. We can increase confidence in
this modeling by continuous validations of the input parameters and its acceptance and usage by
enhancement of its operation.
10
References
Brinkley, J. W. 1985. Acceleration exposure limits for escape system advanced development.
SAE -jornavlo lume 15, No. 2, pp. 10-16.
Brinkley, J. W., and Shaffer, J. T. 1971. Dynamic simulation techniques for the design of escape
systms: Current applications and future Air Force requirements. Wright-Patterson Air Force
Base, Ohio: Aerospace Medical Research Lab. AMRL Technical Report 71-292.
Coltman, J. W. 1986. Crash-resistant crewseat limit-load optimization through dynamic testing,
with a . Simula Inc., Fort Eustis, Virginia: Aviation Applied Technology Directorate,
U.S. Army Aviation Research and Technology Activity, USAAVSCOM TR-85-D-11.
Coltman, J. W., Van Ingen, C., Johnson, N. B., and Zimmermann, R. E. 1989. Aircraft crash
survival design guide. volume 11: Aircraft design crash impact conditions and human
tolerance, Fort Eustis, VA: Aviation Applied Technology Directorate. USAAVSCOM TR
89-D-22B.
Fleck, J. T., Butler, F. E., and Vogel, S. L. 1975. An improved three-dimensional computer
simulation of crash victims. -ITSA reports. DOT-HS-801-507 through 510.
GESAC, Inc. 1992. DYNFEM user's manual. Kearneysville, WV: GESAC, Inc.
Gross, M. E. 1991. The GEBOD III proMgam user's guide and description. Air Force Report ALTR-
1991-0102.
McEntire, B. Joseph. 1992. U.S. Army helicopter inertia reel locking failures. Paper presented at
the annual symposium of the Advisory Group for Aerospace Research and Development
(AGARD), April in Cesme, Turkey.
Obergefell, Louise A., Fleck, John T., Kaleps, Ints, and Gardner, Thomas R. 1988. ArticulatedItoal
body model enhancements: volume 1: Modifications. Wright-Patterson Air Force Base, OH:
Harry G. Armstrong Aerospace Medical Research Laboratory. AAMRL-TR-88-009.
Zimmermann, R. E., and Merritt, N. A. 1989. Aircraft crash survival design guide. Volume I:
Design criteria and checklists. Fort Eustis, VA: Aviation Applied Technology Directorate.
USAAVSCOM TR 89-D-22A.
11
Figure 1. Front and rear seat occupant positions in Apache helicopter.
T = 110 .,. .. ..
T : 0 ,', "' -+..,,
Figure 2. Rear seat occupant motion as conjectured froin accident investigation.
12
3540
.25
Displacement (in)
Figure 3. Helmet profile.
Vo j Helicopter mass
IGravity I Fuselage crush + fracture
Y-+ soil deformation
Tire + oleo
'//I/lIl / // / I ZD/I
Figure 4. Spring-damper-mass idealization of crashing helicopter in a vertical descent.
13
26.2 -
0
._I
Cu
4.4
i'0 It cI ! I
0 12 91 98 133 175
Time (ms)
Figure 5. Vertical (Z) acceleration vs. time for floor of rear seat.
47
114
Ca
(D
0
04.
o 12 98 138
Time (ins)
Figure 6. Vertical (Z) acceleration vs. time for floor of front seat.
14
3540
0
ILLI
.25
Displacement (in)
mom Ws
Figure 7. Seat energy absorber force-displacement characteristic.
0m s 10 ms 1444 ms
* a
Figure 8. Graphics of 180 ms of whole body motion for scenario I simulation.
15
"16 8 ms
Figure 9. Graphics of head and neck motion through glare shield impact for simulation
of scenario 1.
300
.0
0
C -30 -
0
-600-
-900
I I tI I
0 40 80 120 160 200
Time (ms)
MU~ODIOII
Figure 10. Forward (X) component of force in neck vs. time for simulation 1.
16
400
0
-400
0-
o -1200
-1600, I
0 40 80 120 160 200
Time (ms)
MWel Die
Figure 11. Vertical (Z) component of force in neck vs. time for simulation 1.
17
20
0 4
-12
O -28
0
"4)44
0 40 80 120 160 200
Time (ms)
, 21D ie
Figure 12a. Forward (X) component of upper torso acceleration vs. time for simulation 1.
Harness belt acceleration vs. time
Average
13
11
9-. a, 0 7
S3-
<1
-1
-3-
o 20 40 60 80 100 120 140
Time (ms) mnMalaiI
Figure 12b. Harness belt acceleration obtained by double differentiation of output
of simulation 1.
18
1140 ms 180 MS
0
230 ms ._. 250 ms
Figure 13. Graphics of whole body motion through glare shield impact for simulation 2.
" 0MS 80 ms
a a
S~m2s00
Figure 14. Graphics of 200 ms of whole-body motion for simulation 3.
19
10
8
E
00)
CO
CL
541
2
d:0 0 40 80 120 160 200
Time (ms)
MJUODI5O U
Figure 15. Vertical stroke vs. time for front seat in simulation 3.
o 20
4-0
(0
20 408 2 6 0
0
77D
020
20 I
(D-40
600 40 80 120 160 200
Time (ins)
Figure 16. Vertical acceleration vs. time for front seat in simulation 3.
20
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Vicksburg, MS 39180-6199 Sixth U.S. Army
ATTN: SMA
Commanding Officer Presidio of San Francisco, CA 94129
Naval Biodynamics Laboratory
P.O. Box 24907 Commander
New Orleans, LA 70189-0407 U.S. Army Aeromedical Center
Fort Rucker, AL 36362
Assistant Commandant
U.S. Army Field Artillery School Strughold Aeromedical Library
ATTN: Morris Swott Technical Library Document Service Section
Fort Sill, OK 73503-0312 2511 Kennedy Circle
Brooks Air Force Base, TX 78235-5122
Mr. Peter Seib
Human Engineering Crew Station Dr. Diane Damos
Box 266 Department of Human Factors
Westland Helicopters Limited ISSM, USC
Yeovil, Somerset BA20 2YB UK Los Angeles, CA 90089-0021
U.S. Army Dugway Proving Ground U.S. Army White Sands
Technical Library, Building 5330 Missile Range
Dugway, UT 84022 ATTN: STEWS-IM-ST
White Sands Missile Range, NM 88002
U.S. Army Yuma Proving Ground
Technical Library U.S. Army Aviation Engineering
Yuma, AZ 85364 Flight Activity
A'ITN: SAVTE-M (Tech Lib) Stop 217
AFFTC Technical Library Edwards Air Force Base, CA 93523-5000
6510 TW/TTL
Edwards Air Force Base, Ms. Sandra G. Hart
CA 93523-5000 Ames Research Center
MS 262-3
Commander Moffett Field, CA 94035
Code 3431
Naval Weapons Center Commander
China Lake, CA 93555 USAMRMC
ATIN: SGRD-UMZ
Fort Detrick, Frederick, MD 21702-5009
24
Commander Directorate of Training Development
U.S. Army Health Services Command Building 502
ATITN: HSOP-SO Fort Rucker, AL 36362
Fort Sam Houston, TX 78234-6000
Chief
U. S. Army Research Institute USAHEL/USAAVNC Field Office
Aviation R&D Activity P.O. Box 716
ATIN: PERI-IR Fort Rucker, AL 36362-5349
Fort Rucker, AL 36362
Commander, U.S. Army Aviation Center
Commander and Fort Rucker
U.S. Army Safety Center ATTN: ATZQ-CG
Fort Rucker, AL 36362 Fort Rucker, AL 36362
U.S. Army Aircraft Development Chief
Test Activity Test & Evaluation Coordinating Board
ATTN: STEBG-MP-P Cairns Army Air Field
Cairns Army Air Field Fort Rucker, AL 36362
Fort Rucker, AL 36362
Canadian Army liaison Office
Commander Building 602
USAMRMC Fort Rucker, AL 36362
ATrN: SGRD-PLC (COL R. Gifford)
Fort Detrick, Frederick, MD 21702 German Army liaison Office
Building 602
TRADOC Aviation LO Fort Rucker, AL 36362
Unit 21551, Box A-209-A
APO AE 09777 French Army Liaison Office
USAAVNC (Building 602)
Netherlands Army liaison Office Fort Rucker, AL 36362-5021
Building 602
Fort Rucker, AL 36362 Australian Army liaison Office
Building 602
British Army liaison Office Fort Rucker, AL 36362
Building 602
Fort Rucker, AL 36362 Dr. Garrison Rapmund
6 Burning Tree Court
Italian Army liaison Office Bethesda, MD 20817
Building 602
Fort Rucker, AL 36362 Commandant, Royal Air Force
Institute of Aviation Medicine
Farnborough, Hampshire GU14 6SZ UK
25
Defense Technical Information U.S. Army Research and Technology
Cameron Station, Building 5 Laboratories (AVSCOM)
Alexandra, VA 22304-6145 Propulsion Laboratory MS 302-2
NASA Lewis Research Center
Commander, U.S. Army Foreign Science Cleveland, OH 44135
and Technology Center
AIFRTA (Davis) Commander
220 7th Street, NE USAMRMC
Charlottesville, VA 22901-5396 ATTN: SGRD-ZC (COL John F. Glenn)
Fort Detrick, Frederick, MD 21702-5012
Commander
Applied Technology Laboratory Dr. Eugene S. Channing
USARTL-ATCOM 166 Baughman's Lane
ATIN: library, Building 401 Frederick, MD 21702-4083
Fort Eustis, VA 23604
U.S. Army Medical Department
Commander, U.S. Air Force and School
Development Test Center USAMRDALC Liaison
101 West D Avenue, Suite 117 ATTN: HSMC-FR
Eglin Air Force Base, FL 32542-5495 Fort Sam Houston, TX 78234
Aviation Medicine Clinic NVESD
TMC #22, SAAF AMSEL-RD-NV-ASID-PST
Fort Bragg, NC 28305 (Attn: Trang Bui)
10221 Burbeck Road
Dr. H. Dix Christensen Fort Belvior, VA 22060-5806
Bio-Medical Science Building, Room 753
Post Office Box 26901 CA Av Med
Oklahoma City, OK 73190 HQ DAAC
Middle Wallop
Commander, U.S. Army Missile Stockbridge, Hants S020 8DY UK
Command
Redstone Scientific Information Center Dr. Christine Schlichting
ATIN: AMSMI-RD-CS-R Behavioral Sciences Department
/ILL Documents Box 900, NAVUBASE NLON
Redstone Arsenal, AL 35898 Groton, CT 06349-5900
Aerospace Medicine Team Commander
HQ ACC/SGST3 Aviation Applied Technology Directorate
162 Dodd Boulevard, Suite 100 ATrN: AMSAT-R-TV
Langley Air Force Base, Fort Eustis, VA 23604-5577
VA 23665-1995
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Anelor said:

LathspeLL said:

BF3'te çok görüyoruz bunları!

tam bunu dicektim böyle salaklıklar sadece bf3 de oluyo sanıyodum dicektim ahah


bf3 ne kadar gerçekçi oyun anla işte. gerçek hayattaki malları bile alıp koymuşlar oyuna. bunu izleyene kadar asla bilemeyecektik heeh

o değil de şimdi bi daha izlerken fark ettim. yerden sektikten sonra helikopter uçmaya devam ediyor lan hahahaha ama ne uçma
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