October 2014 

NASA/TM–2014-218543 

Evaluation of the First Transport Rotorcraft 

Airframe Crash Testbed (TRACT 1) Full-Scale 

Crash Test 

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October 2014 

NASA/TM–2014-218543 

Evaluation of the First Transport Rotorcraft 

Airframe Crash Testbed (TRACT 1) Full-Scale 

Crash Test 

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ABSTRACT

In 2012, the NASA Rotary Wing Crashworthiness Program initiated the Transport Rotorcraft 

Airframe Crash Testbed (TRACT) research program by obtaining two CH-46E helicopters from 

the Navy CH-46E Program Office (PMA-226) at the Navy Flight Readiness Center in Cherry 

Point, North Carolina. Full-scale crash tests were planned to assess dynamic responses of 

transport-category rotorcraft under combined horizontal and vertical impact loading. The first 

crash test (TRACT 1) was performed at NASA Langley Research Center’s Landing and Impact 

Research Facility (LandIR), which enables the study of critical interactions between the airframe, 

seat, and occupant during a controlled crash environment. The CH-46E fuselage is categorized as 

a medium-lift rotorcraft with fuselage dimensions comparable to a regional jet or business jet. The 

first TRACT test (TRACT 1) was conducted in August 2013. The primary objectives for TRACT 

1 were to: (1) assess improvements to occupant loads and displacement with the use of crashworthy 

features such as pre-tensioning active restraints and energy absorbing seats, (2) develop novel 

techniques for photogrammetric data acquisition to measure occupant and airframe kinematics, 

and (3) provide baseline data for future comparison with a retrofitted airframe configuration. Crash 

test conditions for TRACT 1 were 33-ft/s forward and 25-ft/s vertical combined velocity onto soft 

soil, which represent a severe, but potentially survivable impact scenario. The extraordinary value 

of the TRACT 1 test was reflected by the breadth of meaningful experiments. A total of 8 unique 

experiments were conducted to evaluate ATD responses, seat and restraint performance, cargo 

restraint effectiveness, patient litter behavior, and photogrammetric techniques. A combination of 

Hybrid II, Hybrid III, and ES-2 Anthropomorphic Test Devices (ATDs) were placed in forward 

and side facing seats and occupant results were compared against injury criteria. Loads from ATDs 

in energy absorbing seats and restraints were within injury limits. Severe injury was likely for 

ATDs in forward facing passenger seats, legacy troop bench seats, and a three-tiered patient litter. 

In addition, two standing ATDs were used to evaluate the benefit of Mobile Aircrew Restraint 

Systems (MARS) versus a standard gunner’s belt. The ATD with the MARS survived the impact, 

while fatal head blunt trauma occurred for the standing ATD held by the legacy gunner’s belt. In 

addition to occupant loading, the structural response of the airframe was assessed based on 

accelerometers located throughout the airframe and using three-dimensional photogrammetric 

techniques. Analysis of the photogrammetric data indicated regions of maximum deflection and 

permanent deformation. 

BACKGROUND 

The purpose of NASA’s Fundamental Aeronautics Program Rotary Wing Project (FAP RW) is to 

“develop and validate tools, technologies and concepts to overcome key barriers for rotary wing 

vehicles.” One focal area of FAP RW research is to “foster, develop and demonstrate technologies 

that are critical to the successful commercial operation and passenger acceptance of large rotary 

wing transports in NextGen airspace” (ref. 1). Critical to the acceptance and certification of 

NextGen rotorcraft is the ability to accurately characterize the airframe impact dynamics and 

imparted occupant loads under crash conditions. 

The military crash safety standard for rotorcraft (MIL-STD-1290(A)) specifies occupant seat 

acceleration limits and occupied volume reduction constraints for seven crash impact design 

1
scenarios. These design scenarios are intended to encompass all weight classes and account for 

two impact surfaces, rigid and plowed soil (ref. 2). Crashworthy features, such as energy absorbing 

seats, airbags, and crushable structures, have been designed into rotorcraft such as the AH-64 and 

UH-60 based on this standard. These design features have resulted in an increase in the 90th

percentile survivable longitudinal and vertical impact speeds (ref. 3). A 20-year mishap survey of 

US Navy aircraft by Kent concluded that non-pilot personnel remain more susceptible to injury 

and fatality, and current and future crash protection system improvement should address that 

concern (ref. 4). 

A recent DOD study by Couch and Lindell recommended updates to MIL-STD-1290(A) to include 

a wider range of aircraft classes and impact environments (ref. 5). Recently, the Army Aviation 

Technology Directorate (AATD) sponsored a cooperative effort resulting in the publication of the 

Full Spectrum Crashworthiness Criteria (ref. 6). The FSC outlined guidelines for developing 

comprehensive crash design requirements for a wide range of rotorcraft classes, types, 

configurations, and operating conditions throughout the life cycle of the rotorcraft. The 

Crashworthiness Index (CI) has been proposed as new design standard to replace ADS-11B 

specifications, with a higher score contribution due to basic airframe crashworthiness (ref. 7). 

The Federal Airworthiness Standard for transport category rotorcraft (§14 CFR 29.562) does not 

address crashworthiness at the airframe level (ref. 8). A pair of idealized acceleration vs. time 

conditions is specified for evaluation of the seat and occupant. Crash sled testing with 

Anthropomorphic Test Devices (ATDs) must be conducted to determine seat structural adequacy 

and occupant survivability. Civil seats are often designed with energy absorbers and other 

collapsible mechanisms to limit the occupant loads. The Federal Aviation Administration (FAA) 

is currently developing new policy directives that will address applying computational methods in 

lieu of testing for seat certification (ref. 9). New injury criteria are also being developed by the 

FAA for both side and oblique facing seats (ref. 10). 

The performance of crashworthy systems is optimized primarily for a specific occupant weight 

(typically 50th percentile males) and assumed idealized input deceleration pulses. However, the 

response is highly sensitive to any variations from those standards. Multi-terrain impact (water, 

soft soil, prepared surface) will alter the magnitude and duration of the airframe deceleration. The 

combination of landing gear stroke, subfloor crushing, floor and frame deformation, seat stroke,

and restraint activation must all be taken in account. 

Essential in this effort is the development and assessment of modeling tools that can accurately 

associate impact velocities, attitudes and terrains to seat interface and occupant G-loads. Recent 

advancements in computational techniques have allowed for streamlined and efficient evaluations 

of the crash performance of rotorcraft. Finite element models (FEM) have been developed that 

contain sufficient fidelity to model material damage and failure progression within the airframe 

during impact and are yet computationally affordable. The analytical capabilities have progressed 

from vehicle lumped parameter models (KRASH) to multi-body occupant models (MADYMO) to 

high fidelity explicit finite element analyses (LS-DYNA, PAM-CRASH, Radioss, ABAQUSExplicit). Detailed representations of seats, occupants, and restraints can be feasibly included in a 

system-level FEM to accurately account for the load transfer between the airframe and the 

occupant and to directly assess the likelihood of occupant injury. 

2
The current emphasis in aviation across all categories is incorporation of high performance 

composite materials into the airframe design to enhance vehicle efficiency. Examples of medium 

to heavy lift rotorcraft containing more than 20 percent composites by weight include the V-22 

Osprey and the CH-53K. The benefit of increased stiffness to weight ratio provided by composites 

must be balanced against the crashworthiness of the structure. Composite materials have unique 

considerations for both performance evaluation and certification. The impact dynamics behavior 

of composites is markedly different than legacy aluminum airframes in terms of failure initiation, 

damage progression, and specific energy absorption. Analytical tools and material/structural 

characterization tests are necessary for the design of energy-absorbing composite structures. The 

current design building block test sequence for material characterization and analysis validation 

methodologies may not encapsulate all critical modes of failure. Current models are 

phenomenological and parameters in simulations are determined by calibration from test. 

Ultimately, crash safety certification by analysis is sought to minimize reliance on full-scale crash 

tests. Nonetheless, a requisite number of section-level to full-scale tests must be conducted to 

anchor simulations of rotorcraft which incorporate novel crashworthy features and composite 

airframe structures. With the support of the FAP RW Project, two full-scale crash tests were 

planned to address transport category rotorcraft crash responses under combined horizontal and 

vertical loading. The Transport Rotorcraft Airframe Crash Testbed (TRACT) full-scale crash tests 

were proposed. The first test, designated TRACT 1, would evaluate a collection of crashworthy 

system experiments and characterize the metallic fuselage response to a severe but survivable 

impact. The second test (TRACT 2) will contain similar experiments but also include a retrofit of 

the TRACT subfloor with composite energy absorbing concepts. This memorandum is intended to 

provide specific details regarding TRACT 1. 

INTRODUCTION 

Test Facility 

A photograph of NASA Langley’s Landing and Impact Research Facility (LandIR) is shown in

Figure 1. The LandIR facility provides the unique capability for full-scale impact testing with 

combined vertical and longitudinal loading onto prepared surfaces, soil, and water. Combined 

loading tests are conducted by suspending the test article from the gantry structure using two sets 

of cables: pullback cables and swing cables. These cables are attached to the airframe at hard points 

that enable the helicopter to be lifted through its center of gravity. The airframe is lifted using the 

pullback cables to a specified height and pyrotechnically released following a countdown. Swing 

cables are configured to form a parallelogram to minimize pitch angular velocity during the 

pendulum swing prior to impact. For land impact, just prior to ground contact, the supporting 

cables are pyrotechnically separated. For water impact, the test article is suspended from a carrier 

platform which remains attached to the swing cables. The test article is pyrotechnically severed 

from the platform and freely impacts the Hydro-Impact Basin. 

The LandIR has 320-channels of data acquisition (DAS) capability to record accelerations, strains, 

displacements, and forces. Three-dimensional photogrammetry is recorded with as many as 1,000 

3
frames per second and synchronized to DAS results using an IRIG time code generator. External 

vehicle kinematics and onboard occupant responses are tracked using high speed and high 

definition ruggedized cameras. Eleven Hybrid II and III ATDs are available ranging in size from 

3 year and 6 year old child to 5th to 95th percentile adult. 

Figure 1. Landing and Impact Research Facility (LandIR) 

Test Article 

The CH-46E Sea Knight is categorized as a medium-lift tandem rotor helicopter with length and 

width of nearly 45 ft. and 7 ft., respectively, and a capacity for 5 crew and 25 troops. A schematic 

drawing is shown in Figure 2 showing Fuselage Station (FS), Water Line (WL), and Butt Line 

(BL) locations. These dimensions are given in units of inches. The fuselage design is 

semi-monocoque with skin stiffeners and frame sections. The cabin fuselage cross section is nearly 

uniform and is composed primarily of Aluminum 2024 and 7075 alloys. Fuselage frames are 

spaced longitudinally approximately every 30 inches. Between FS120 and FS410, the fuselage 

section has a nearly uniform cross-section. The FS locations of major structural features are listed 

in Table 1.

240’

400’

265’

Hydro-Impact 

Basin 

constructed 

2010

(a) 70-foot Drop Tower (b) Gantry

4
Figure 2. CH-46E airframe schematic with FS, WL, and BL locations highlighted. 

Table 1. Longitudinal FS Locations of major structural features 

Feature FS Location

Cockpit and cockpit bulkhead FS 0 - FS 120

Main cabin FS 120 – FS 320

Stubwing/aft cabin FS 320 – FS 410

Tail FS 410 – FS 490

In 2012, NASA Langley obtained two CH-46E fuselages from the Navy CH-46E Program Office 

(PMA-226) that were located at the Navy Flight Readiness Center/Cherry Point. The stubwings 

which hold the main landing gear were removed for testing, along with the vertical tail and engine, 

rotor, and transmission components. The weight of these components was not replaced, because 

the fuselage was intended not to mimic the overall weight and center of gravity (CG) of the CH46E. The main cabin from the cockpit aft bulkhead to the stubwing area contains five frame 

sections that are nearly identical in spacing and design to allow multiple occupant, seat, and 

crashworthy systems experiments to be evaluated. The as-delivered bare fuselage weight was 

approximately 2,500 lb. A photograph of the TRACT 1 airframe is shown in Figure 3. 

.

5
Figure 3. TRACT 1 fuselage 

Modifications were made to the TRACT structure before installation of the experiments. The 

aluminum honeycomb floor was reinforced underneath to provide hard points for mounting seats 

and other experiments. Two swing beams were bolted to the sidewalls approximately 3 inches 

above the waterline to provide a sturdy interface for the swing and pullback cables. Fiberglass 

panels were fabricated and attached to the underside of the cockpit enclosure to reduce plowing of 

the exposed cockpit. 

Collaborators 

The primary objective for TRACT 1 was to evaluate the integrated airframe, seat, and occupant 

responses under a combined horizontal and vertical impact velocity. This test article provided an 

exceptional opportunity for numerous onboard experiments, seat configurations, and ATDs to be 

included and evaluated. Discussions with crashworthiness and injury biomechanics organizations 

within the Federal Aviation Administration (FAA), Department of Defense (DOD), and rotorcraft 

industry manufacturers led to formal agreements with four collaborators: 

x The Naval Air Systems Command (NAVAIR), Human Systems Department, Crashworthy 

Systems Branch, Patuxent River, MD, develops, evaluates, and qualifies systems for Naval 

Aviation intended to prevent injury resulting from impact-based aviation mishaps. 

x The FAA Civil Aeromedical Institute (CAMI), Aeromedical Research Division, Protection and 

Survival Laboratory, Biodynamics Research Team, conducts research concerning occupant 

impact protection in civil aircraft. To evaluate the protection provided by seats and restraint 

systems, the team develops new testing protocols, test dummy modifications and new injury 

criteria. 

x The U.S. Army Aeromedical Research Laboratory (USAARL), Warfighter Protection 

Division, Injury Biomechanics Branch, investigates air and ground warfighter response to 

dynamic loading, including blast, ballistics, and impact. 

x Cobham Life Support develops restraint systems for fixed and rotary wing and ground 

vehicles. 

6
Experiments 

Test Objective 1: Comparison of the CH-46 crew seat with MA-16 inertia reel and CH-46 crew 

seat with PARS. 

The CH-46E crew seat contains two pairs of energy absorbers (EAs). One pair of variable-load 

EAs, aligned aft of the seat back, is designed to stroke and limit the loads along the spinal direction. 

A second pair of fore/aft-oriented EAs limits the seat rail interface loads and prevents detachment 

of the seat during forward deceleration and occupant flail. The pilot seat is shown in Figure 4. 

Figure 4. CH-46E crew seat

Five-point restraints are used with the CH-46E seat. The shoulder webbings are spooled onto the 

MA-16 inertia reel, which locks passively under inertial loads. As much as 5 inches of extra 

webbing payout is possible due to pack-down on the spool and elongation due to flail, or forward 

body displacement. The Pre-Tensioning Aircrew Restraint System (PARS) system is intended to 

decrease the payout through pyrotechnically actuated spool retraction. Cobham Life Support 

provided the MA-16 reels, and the PARS pretensioning devices and control modules. The PARS 

system can be activated with customary crash sensors. For the TRACT 1 test, the PARS was 

activated at the point of impact with redundant contact switches mounted on the aft belly region. 

The MA-16 inertia reel and PARS system are located behind the seat back in between the vertical 

EAs, as shown in Figure 5. 

7
Figure 5. MA-16 inertia reel with PARS 

NAVAIR provided two fully instrumented 50th percentile Aerospace Hybrid III ATDs for the pilot 

and co-pilot seats. The ATDs were outfitted with standard flight suits, helmets, and flight gear. 

The test article did not have a cockpit enclosure forward of the directional pedals. Therefore, 

aluminum floor pans were added underneath the feet to ensure loads could be distributed through 

both the legs and the pelvis. The cockpit ATDs are shown in Figure 6. 

Figure 6. Cockpit ATDs 

MA-16

PARS

8
Test Objective 2: Floor-mounted passenger ATD response 

The civilian standard for transport category rotorcraft (14 CFR 29.562) specifies two distinct 

impact conditions that must be tested or analyzed to determine seat adequacy and occupant 

protection (ref. 8). Tests are normally conducted using acceleration or deceleration crash sleds and 

a seated 50th percentile Hybrid II or 50th percentile FAA Hybrid III ATD. The first test combines 

vertical and longitudinal deceleration with a seat pitched 60 degrees relative to the sled velocity 

vector. The impact velocity is 30 feet per second with a rise time of no more than 0.031 seconds, 

resulting in a peak deceleration of 30 g. The second test is mostly longitudinal, with no seat pitch 

and a yaw angle of 10 degrees. The impact velocity is 42 feet per second, with a rise time of no 

more than 0.071 seconds, resulting in a peak deceleration of 18.4 g. Seat compliance is based on 

evaluation of whether the seat detaches from the rails, the seat frame collapses, or the seat rails 

excessively warp. Occupant compliance involves two criteria. If head blunt trauma occurs, the 

Head Injury Criteria (HIC) index, which is computed through integration of the resultant head 

acceleration time history, cannot exceed a value of 1,000. Lumbar compressive loads, measured 

from a load cell below the straight lumbar spine block, cannot be greater than 1,500 lb. For lap 

belt only configurations, no restraint limits are specified. For a three-point shoulder belt, shoulder 

restraint loads cannot exceed 1,750 lb. For a four or five point harness, shoulder restraint loads 

cannot exceed 2,000 lb. For both tests, the input acceleration pulse shapes are triangular. 

There are two predominant mechanisms of energy absorption that occur within the cabin during 

impact, seat frame and cushion distortion, and floor and subfloor web deformation. The timing and 

interaction of these mechanisms, combined with the occupant flail response, influence the resultant 

occupant dynamic loading. It is valuable to compare the response of the seat interface in a fullscale crash test environment against the idealized input pulses prescribed for certification. 

Two pairs of passenger aircraft seating systems certified to Part 25 standards (ref. 11) were donated 

by aircraft seat manufacturers. The first seat pair contained a diagonal strut to absorb loads. The 

second seat pair contained lightweight seat legs without dedicated energy absorbers. Seat legs were 

adjusted and centered to 23 inches of spacing to share a single pair of seat rails. Four LandIR ATDs 

were used for these seats. A 50th percentile Hybrid II and 50th percentile FAA Hybrid III were 

seated in the first pair. A 5th percentile female and a 95th percentile male were seated in the second 

pair. The floor mounted cabin seats with ATDs are shown in Figure 7. The floor at FS 254 was 

also loaded with 400 lb. of ballast mass to provide mass loading on the subfloor comparable to two 

pairs of seats. 

9
Figure 7. Floor-mounted cabin seats with ATDs 

Test Objective 3: Comparison of standing ATD with traditional gunner’s belt and standing ATD 

with Aircrew Endurance Vest and Mobile Aircrew Restraint System (MARS). 

Aircrews working within military helicopters are regularly tethered to various anchor points in the 

cabin to protect against falls during a mishap. This tether is a nylon strap commonly known as a 

gunner’s belt. The gunner’s belt is tied to a chest strap and is typically adjusted to the longest 

length possible for ease of mobility. Mishap data suggests that a slackened gunner’s belt may cause 

injury and fatality during mishaps from strap entanglement, crew impact with structure, or crew 

excursion. A gunner’s belt system that can extend and retract as the crew moves about the cabin 

is preferable to reduce the strike envelope. 

10
The MARS system was developed by Cobham Life Support as a variation on the MA-16 inertia 

reel (ref. 12). The MARS mounted in the TRACT test article is shown in Figure 8. The reel can be 

mounted along multiple cabin ceiling locations, and extends and retracts the webbing as the 

aircrew moves about the cabin. In a mishap, the retractor locks and slack is minimized. The MARS 

has been evaluated at the component and crash sled level under dynamic loads, and in flight tests, 

but not in a full-scale combined velocity impact environment. 

Figure 8. MARS 

NAVAIR provided (2) fully instrumented 5th percentile male Aerospace Hybrid III ATDs, one 

attached to the MARS, and one attached to a standard gunner’s belt. The smaller 5th percentile 

ATDs were used due to cabin height. To secure the standing ATDs in position during pullback and 

swing, the back of each neck was supported by thin steel cabling, and stabilized by parachute 

cords. A single pyrotechnic cable cutter was used to sever the steel cable at the point of impact. 

Hook and loop pads were used to hold boots in place. A photograph showing both standing ATDs 

is shown in Figure 9. 

11
Figure 9. Standing ATDs 

Test Objective 4: Full-field photogrammetry experiment. 

Full-field deformation techniques, which include scanning vibrometers, various types of 

interferometry techniques and digital image correlation, have become more commonplace within 

the past decade. Digital image correlation applications have vastly expanded due to the 

advancements of computer processing power required for pattern recognition algorithms, 

increasing robustness of the sensor hardware, and cost reductions. Novel application ideas have 

allowed these techniques to shift from strain gauge systems in laboratory data acquisition devices 

to established data acquisition techniques in the field. 

Onboard data recorders provide some indication of impact parameters such as speed and angle 

during a mishap. Typically, forensic reconstruction techniques can also piece together what 

happened; however, the actual loads and precise deformation on the airframe are usually unknown.

One advantage of controlled crash testing is the capability to measure the damage progression of 

the airframe. The full-field deformation imaging technique was used in this instance to determine, 

for the first time, quantitatively, the overall deformation characteristics of a large vehicle airframe 

during a realistic impact event. 

Two monochrome high speed cameras viewed and recorded images of the impact from slightly 

different angles at a rate of 500 frames per second resulting in pairs of synchronized images of the 

12
impact. A random pattern of over 8,000 1-inch diameter dots was applied to the port airframe skin 

from the cockpit to the tail. After a camera calibration procedure is completed, the cameras are 

able to determine the position of each airframe dot in space. Pattern recognition algorithms sense 

relative deformation between dot groups and an entire deformation field is constructed. The dot 

pattern for the TRACT and an example of a fringe plot overlay is shown in Figure 10. During test 

preparation, the airframe was loaded onto a trailer and pulled across the gantry to calibrate the 

photogrammetry system. The colors in Figure 10 represent relative distance traveled based on a 

fixed camera location. 

Figure 10. Full-field photogrammetry, strain pattern 

Test Objective 5: Comparisons of sidewall-mounted Crew Attenuating Crew Seat (CACS) troop 

seated Hybrid III ATD, seated Hybrid III ATD with ES-2re head and neck, and two seated 

Hybrid II ATDs in a standard CH-46 tube and rag sidewall troop seat. 

The FAA dynamic test requirements described for Experiment 2 were specifically intended to 

address forward and aft facing seats. FAA research has shown that passengers in side-facing seats 

are exposed to injury risks from head and upper extremity flail and pelvic excursion that are not 

observed for passengers in forward or aft-facing seats (ref. 10). To address the safety of side-facing 

seats in transport category aircraft, the FAA provided policy guidance that adopted portions of the 

automotive side-facing requirements, including use of the ES-2re ATD. The ES-2re is designed to 

evaluate the injury potential of contact and inertial loads produced during a side impact. Since the 

automotive side-facing standards do not address neck loading, it was necessary for the FAA to 

sponsor development of new upper neck injury criteria. Much of the FAA side facing seat research 

was conducted on crash sleds that introduce lateral loads. The TRACT test provides the 

opportunity to include both lateral and vertical loading in a realistic impact scenario. 

The CH-46E troop seats are side-facing. The legacy CH-46E troop bench is an aluminum seat pan 

frame with canvas mesh overwrap and lap restraints. The CH-46E Crew Attenuating Crew seat 

(CACS) is an energy absorbing foldable seat with a five-point restraint. Two wire bender struts 

provide vertical load limiting capability for the restrained occupant. The typical troop bench and 

CACS configurations are illustrated in Figure 11. 

13
Figure 11. CH-46E troop seats 

CAMI provided a fully instrumented FAA 50th percentile Hybrid III ATD with an ES-2re head 

and neck for use on one CACS seat. The FAA Hybrid III construction differs from the standard 

model by including components to permit accurate evaluation of injury risks due to vertical 

(through the spine) loads. The ES-2re head and neck components provide more biofidelic 

head/neck kinematics in the lateral direction and injury criteria are available to relate the neck 

loads measured to injury risk. NAVAIR provided a fully instrumented 50th Aerospace Hybrid III 

ATD in a second CACS seat for comparison to the CAMI ATD. The Aerospace Hybrid III can 

also evaluate injury risk from vertical loading. It uses a “pedestrian” pelvis assembly that provides 

more freedom of motion for the lower legs than provided by the “sitting” pelvis design of the FAA 

Hybrid III. Both Hybrid III models have shoulder and hip construction that provide consistent belt 

interaction. Figure 12 shows the CACS side-facing ATDs. Two 50th percentile Hybrid II LandIR 

ATDs were seated on the troop bench and restrained with lap belts (Figure 13). It was anticipated 

the troop bench would not survive the impact undamaged. Therefore, the troop bench experiment 

functioned mostly as a performance benchmark when evaluating the response of side-facing 

occupants. 

14
Figure 12. CACS side-facing ATDs 

Figure 13. Troop bench ATDs 

15
Test Objective 6: Evaluation of markerless tracking on ATD response. 

Traditional tracking systems use targets, stickers, paint or other externally applied features which 

discriminate specific areas (or targets) on the object. Target motion is captured with high speed 

cameras to measure kinematic motion and flail envelope. The development in markerless tracking 

capabilities and systems could potentially change the way measurements are made on ATDs. 

Markerless tracking systems use their own projection/acquisition system to acquire data on either 

specific features or on the entire object without the aid of external markers. These measurement 

systems are still in development, but some of the early systems show promise as an alternative or 

supplement to either the conventional tracking systems or the internal sensor suite. 

An experiment to utilize commercial markerless tracking hardware within a dynamic impact 

environment was proposed. The standing ATD with the MARS restraint provided a suitable target 

for tracking. The commercial hardware consisted of a color digital video camera, along with a 

depth sensor/camera and an infrared projector combined in one package. The depth sensor and 

infrared projector worked in conjunction to obtain depth measurements in the field of view via a 

projection/comparison system. The infrared projector projected an infrared dot grid into the field 

of view, while the depth sensor measured the distance between all of the individual projected dots. 

The sensor then compared the calculated distances to a calibrated value. The difference between 

these two values was the measured depth of the projected dots under interrogation. When measured 

over the entire field of view, a full field depth field was generated. Internal logic preprogrammed 

with the sensor further examined the depth field and determined the presence of a human by 

employing various edge detection and area centroid algorithms. If a human shape was detected, a 

“best fit” three-dimensional skeletal wireframe could then be computed. The wireframe data can 

be manipulated to give three-dimensional positional time histories of 19 individual joint locations 

present in the body. From time histories of the joint locations, joint forces could be related to 

conventional ATD injury criteria. Figure 14 shows the camera field of view and depth fringe plot. 

Figure 14. Markerless Tracking 

16
Test Objective 7: Comparison of cargo experiment with non-energy-absorbing restrained cargo 

mass and energy absorbing restraint cargo mass. 

Military and civilian crashworthiness standards require retention of high mass items such as 

equipment, luggage, or cargo during a mishap. High mass cargo is often restrained to cargo holds 

and side walls with netting or straps. As the cargo reacts against the restraints, the loads transmitted 

at the restraint supports can exceed the structural capability of the interface, and the restraints can 

break free. An optimal cargo restraint design is robust enough to withstand the deceleration of the 

cargo without structural failure at the interface. 

A load-limiting cargo restraint was developed by Penn State University that uses a stitch ripping 

device (SRD). Energy is absorbed by webbing extension, thread rupture and stitch slippage. The 

restraint was attached to a breakaway fitting rated for 5,000 lb. failure load. The fitting was then 

connected to a 500 lb. sliding mass. A standard nylon webbing restraint was also evaluated 

alongside the SRD restraint with the same breakaway fitting and sliding mass. With 10 g

deceleration, the loads would be exceeded in the standard restraint, causing detachment of the 

mass. The load-limiting restraint was designed to limit to 3,000 lb., and would therefore stroke 

without fitting detachment. The load-limiting experiment is shown in Figure 15. It was fabricated 

to be self-contained on a platform that could be bolted to the aft TRACT cabin floor. Housings 

were assembled around the masses for containment during deceleration. 

Figure 15. Cargo Restraint Experiment 

17
Test Objective 8: Three-tiered litter mounted in the aft starboard region of the aircraft cabin. 

Legacy litter systems in military rotorcraft have been qualified under static loading. These 

standards have not been updated in the same manner as have crashworthy seats. An understanding 

of the performance of the litters and supine litter patients in a crash-induced dynamic loading event 

is not well understood. 

The litter experiment involved a combination of DOD resources. USAARL supplied a three-tiered 

litter stanchion system compatible with the CH-46. The stanchion interfaces with existing sidewall, 

ceiling, and cargo rail supports. Two litters, one instrumented 50th Aerospace Hybrid III ATD, and 

one 200 lb. Grumman-Alderson Research Dummy (GARD) manikin were provided by the 

USAARL. NAVAIR provided one litter and a second GARD manikin. The two GARD manikins 

provided the proper mass loading of the litter, which is depicted in Figure 16.

Figure 16. Patient litter experiment 

Instrumentation Plan 

For TRACT 1, 353-channels of data were collected at 10,000 samples per second using onboard 

digital data acquisition systems. Schematics of the experiments, instrumentation layout, and 

cameras are provided in Figures 17-23. In addition, the channel number and sensor types for each 

airframe location or experiment are listed in Table 2. Tables 3-6 list the individual ATD 

instrumentation channels. The overall distribution of instrumentation is: 

• 279 channels devoted to internal instrumentation of the ATDs and restraint load cells 

18
• 61 channels designated for accelerometers that were mounted on blocks at stiff interfaces 

between frames and skin, or on ballast weight 

• 6 channels for lifting cable load cells 

• 7 IRIG time code channels for synchronization with high speed cameras 

CAMI provided a DAS rack that linked with the LandIR DAS system and would record all 

responses for their ATD. USAARL provided DAS racks that were also linked with the LandIR 

DAS system and would record litter instrumentation along with other sensors located in the aft 

cabin. The LandIR, CAMI, and USAARL DAS racks were connected in series to receive common 

trigger, onboard battery power, and DAS software control. NAVAIR supplied a separate DAS that 

was compatible with the NAVAIR ATD instrumentation. The NAVAIR DAS was independently 

controlled and powered with onboard batteries. The DAS racks were installed at two locations on 

the cockpit bulkhead and on a sidewall-mounted shelf aft of the second forward facing seat pair. 

Figure 17. Cockpit instrumentation 

1

2

11

12

13

ATD Location

Accelerometer

Location

High speed camera

X

Z

Y

19
Figure 18. Cabin instrumentation- ATD and high speed camera, isometric view 

Figure 19. Cabin instrumentation- accelerometers 

3 4

5 6

7

8

9

10

ATD/Experiment Location

High speed camera

32

31

X

Z

Y

33

34

15

16

14

17

18

19

20

21

22

23

24

25

26

27

29

30

28

Accelerometer Location

X

Z

Y

20
Table 2. Data acquisition instrumentation list 

Position Measurement Location Notes Total # of 

Channels

1 ATD Aerospace Hybrid 

III 50th male (NAVAIR) Co-pilot (front left)

Accelerometers, spinal load 

cells, shoulder strap load 

cells

28

2 ATD Aerospace Hybrid 

III 50th male (NAVAIR) Pilot (front right)

Accelerometers, spinal load 

cells, shoulder strap load 

cells

28

3 ATD Hybrid III/ES-2

Neck (CAMI)

Cabin (side-facing CACS

Seat)

Accelerometers, spinal load 

cells, shoulder strap load 

cells

51

4 ATD Aerospace Hybrid 

III 50th male (NAVAIR)

Cabin (side-facing CACS

Seat)

Accelerometers, spinal load 

cells, shoulder strap load 

cells

28

5 ATD Hybrid II 50th

male (NASA LandIR)

Cabin (side-facing troop 

bench)

Accelerometers, spinal load 

cells, lap strap load cells 11

6 ATD Hybrid II 50th

male (NASA LandIR)

Cabin (side-facing troop 

bench)

Accelerometers, spinal load 

cells, lap strap load cells 11

7 ATD Hybrid II 50th

male (NASA LandIR)

Cabin (forward facing 

passenger seat)

Accelerometers, spinal load 

cells, lap strap load cells 11

8

ATD FAA Hybrid III 

50th male (NASA

LandIR)

Cabin (forward facing 

passenger seat)

Accelerometers, spinal load 

cells, lap strap load cells 17

9 ATD Hybrid III 95th

male (NASA LandIR)

Cabin (forward facing 

passenger seat)

Accelerometers, spinal load 

cells, lap strap load cells 17

10 ATD Hybrid III 5th

female (NASA LandIR)

Cabin (forward facing 

passenger seat)

Accelerometers, spinal load 

cells, lap strap load cells 17

11 Acceleration Co-pilot (front left floor) X, Z 2

12 Acceleration Between pilot-co-pilot X, Y, Z 3

13 Acceleration Pilot (front right floor) X, Z 2

14 Acceleration Main Rotor X, Y, Z 3

15 Acceleration

Cockpit/cabin fuselage 

splice, left floor (FS 

160.0)

X, Y, Z 3

16 Acceleration

Cockpit/cabin fuselage 

splice, right floor (FS 

160.0)

X, Y, Z 3

17 Acceleration Cabin Fuselage, left floor 

(FS 220.0) X, Z 2

18 Acceleration Cabin Fuselage, right 

floor (FS 220.0) X, Z 2

19 Acceleration Cabin Fuselage, right 

side wall (FS 220.0) X, Z 2

20 Acceleration Cabin Fuselage, left floor 

(FS 254.0) X, Z 2

21 Acceleration Cabin Fuselage, right 

floor (FS 254.0) X, Z 2

21
Position Measurement Location Notes Total # of 

Channels

22 Acceleration Cabin Fuselage, right 

side wall (FS 254.0) X, Z 2

23 Acceleration Cabin Fuselage, left floor 

(FS 286.0) X, Z 2

24 Acceleration Cabin Fuselage, right 

floor (FS 286.0) X, Z 2

25 Acceleration Cabin Fuselage, right 

side wall (FS 286.0) X, Z 2

26 Acceleration Cabin Fuselage, left floor 

(FS 350.0) X, Y, Z 3

27 Acceleration Cabin Fuselage, right 

floor (FS 350.0) X, Y, Z 3

28 Acceleration Aft Rotor X, Y, Z 3

29 Acceleration Cabin Fuselage, left floor 

(FS 410.0) X, Z 2

30 Acceleration Cabin Fuselage, right 

floor (FS 410.0) X, Z 2

31 Pedestrian ATD (5th 

HIII) Aft of forward bulkhead Accels, Gunner’s belt webbing 10

32 Pedestrian ATD (5th 

HIII) Aft of forward bulkhead Accels, MARS webbing 10

33 Litter ATD Aft Cabin (near FS 

382.0)

Litter interfaces forces and 

accels 51

34 Cargo restraint Aft Cabin (near FS 

382.0)

Accels and load cells on 

masses

11

NA Load cell 4 Swing, 2 Pullback 

cables 10,000-lb capacity 6

22
Figure 20. ATD Coordinate System and instrumentation Location 

Head

Upper Neck

Lower Neck

Pelvis/Lumbar

Thorax

L

Femur

23
Table 3. NAVAIR ATDs 

1 Head Ax 1 Head Ax

2 Head Ay 2 Head Ay

3 Head Az 3 Head Az

4 Upper Neck Fx 4 Upper Neck Fz

5 Upper Neck Fy 5 Upper Neck Mx

6 Upper Neck Fz 6 Upper Neck My

7 Upper Neck Mx 7 Chest Ax

8 Upper Neck My 8 Chest Ay

9 Upper Neck Mz 9 Chest Az

10 Lower Neck Fx 10 Chest Deflection

11 Lower Neck Fy

12 Lower Neck Fz

13 Lower Neck Mx

14 Lower Neck My

15 Lower Neck Mz

16 Chest Ax

17 Chest Ay

18 Chest Az

19 Pelvic Ax

20 Pelvic Ay

21 Pelvic Az

22 Lumbar Fx

23 Lumbar Fx

24 Lumbar Fz

25 Lumbar Mx

26 Lumbar My

27 Lumbar Mz

28 Strap Load (on shoulder belt)

NAVAIR Hybrid III 

Channel Count

NAVAIR Pedestrian 5th 

Channel Count

24
Table 4. FAA Hybrid III/ES-2 50th ATD channel count

1 HIII Head Ax 30 HIII Pelvis Ax

2 HIII Head Ay 31 HIII Pelvis Ay

3 HIII Head Az 32 HIII Pelvis Az

4 HIII Head Rx ARS18K 33 HIII Pelvis Rx ARS12K

5 HIII Head Ry ARS18K 34 HIII Pelvis Ry ARS12K

6 HIII Head Rz ARS18K 35 HIII Pelvis Rz ARS12K

7 ES2 Upper Neck Fx 36 HIII Lumbar Fx

8 ES2 Upper Neck Fy 37 HIII Lumbar Fy

9 ES2 Upper Neck Fz 38 HIII Lumbar Fz

10 ES2 Upper Neck Mx 39 HIII Lumbar Mx

11 ES2 Upper Neck My 40 HIII Lumbar My

12 ES2 Upper Neck Mz 41 HIII Lumbar Mz

13 HIII Lower Neck Fx 42 Left Femur Rx ARS18K

14 HIII Lower Neck Fy 43 Left Femur Ry ARS18K

15 HIII Lower Neck Fz 44 Left Femur Rz ARS18K

16 HIII Lower Neck Mx 45 Right Femur Rx ARS18K

17 HIII Lower Neck My 46 Right Femur Ry ARS18K

18 HIII Lower Neck Mz 47 Right Femur Rz ARS18K

19 HIII Thorax Ax 48 Upper Left Shoulder Strap

20 HIII Thorax Ay 49 Upper Right Shoulder Strap

21 HIII Thorax Az 50 Left Lap Belt

22 HIII Thorax Rx ARS12K 51 Right Lap Belt

23 HIII Thorax Ry ARS12K

24 HIII Thorax Rz ARS12K

25 Thoracic Fx

26 Thoracic Fy

27 Thoracic Fz

28 Thoracic Mx

29 Thoracic My

25
Table 5. USAARL ATD channel count 

1 Head AX 31 Right Femur MX 

2 Head AY 32 Right Femur MY

3 Head AZ 33 Right Femur MZ

4 Neck FX 34 Fwd Litter Strap z

5 Neck FY 35 Fwd Litter Strap y

6 Neck FZ 36 Fwd Litter Strap x

7 Neck MX 37 Aft Litter Strap z

8 Neck MY 38 Aft Litter Strap y

9 Neck MZ 39 Aft Litter Strap x

10 Chest AX 40 Fwd Wall Mount z

11 Chest AY 41 Fwd Wall Mount y

12 Chest AZ 42 Fwd Wall Mount x

13 Lumbar Spine FX 43 Aft Wall Mount z

14 Lumbar Spine FY 44 Aft Wall Mount y

15 Lumbar Spine FZ 45 Aft Wall Mount x

16 Lumbar Spine MX 46 Top Fwd Belt

17 Lumbar Spine MY 47 Top Litter Aft Belt

18 Lumbar Spine MZ 48 Mid Fwd Belt

19 Pelvis AX 49 Mid Aft Belt

20 Pelvis AY 50 Btm Litter Fwd

21 Pelvis AZ 51 Btm Litter Aft Belt

22 Left Femur FX 

23 Left Femur FY

24 Left Femur FZ

25 Left Femur MX 

26 Left Femur MY

27 Left Femur MZ

28 Right Femur FX 

29 Right Femur FY

30 Right Femur FZ

26
Table 6. NASA LandIR ATDs 

Figure 21. Cabin instrumentation- ATD and high speed camera, top view 

1 Head Ax 1 Head Ax

2 Head Ay 2 Head Ay

3 Head Az 3 Head Az

4 Upper Neck Fx 4 Chest Ax

5 Upper Neck Fz 5 Chest Ay

6 Upper Neck My 6 Chest Az

7 Lower Neck Fx 7 Lumbar Fz

8 Lower Neck Fz 8 Pelvic Ax

9 Lower Neck My 9 Pelvic Ay

10 Chest Ax 10 Pelvic Az

11 Chest Ay 11

12 Chest Az

13 Lumbar Fz

14 Pelvic Ax

15 Pelvic Ay

16 Pelvic Az

17 Strap Load (on lap belt)

Strap Load (on lap belt)

NASA Hybrid III Channel 

Count (5th,50th,95th)

NASA Hybrid II Channel 

Count

1

2

7

8

3 4 5 6

9 31

32 33

10

34

Redlake 8mm

Miro3 - 8mm Miro ex 8mm 

Miro3 - 8mm 

Kinect Sensor

Redlake BW 1MP 8mm

Miro 3 8mm

Miro ex 12mm 

27
Figure 22. Cabin instrumentation ATD and high definition camera- top view 

Figure 23. External camera layout 

1

2

7

8

3 4 5 6

9 31

32 33

10

34

Hero1

1080p 

Hero 3 cams film 720 @ 120 fps

Hero 1 cams film 1080p @ 60 fps

Hero3

120fps 

Hero3

120fps 

Hero1

1080p 

Hero1

1080p 

Hero3

1080p 

Hero3

120fps

Hero3

120fps

Hero3

1080p

Hero1

1080p

Hero3

1080p

TRACT

Backboard

N

P10 BW

Data L

P10 BW

Data R

P9 50mm

South

P9 24mm

Southeast

ground

P9 24mm

West

P12 BW 12mm

West

P12 BW 12mm

Roof

P9 24mm

Southwest ground

P9 35mm

East

P9 35mm

North

28
Weight and Balance Data 

The locations of the four swing cable and two pullback cable interfaces to the TRACT test article 

were dependent on the location of the CG. Any misalignment of the cables relative to the CG 

would cause the test article to pitch and roll upon release, causing a deviation from the intended 

impact attitude. To verify the predicted weight and balance estimates, the test article was 

suspended from three cables and pitched up to determine the CG in all three axes, as shown in 

Figure 24. The fully-loaded, fully-instrumented test article weighed 10,300-lb. The vertical CG 

was 3 inches above the swing cable and 6 inches above the waterline. The lateral CG was less than 

1 inch from the centerline. The longitudinal CG was near FS 260.0. The actual weight and CG 

location matched well with the estimates. 

Figure 24. Photograph taken during weight and balance testing 

Impact Conditions 

The impact conditions that were defined for the TRACT 1 test are considered to be severe but 

survivable, based on a survey of mishap data. Figure 25 shows a plot of the 95th percentile velocity 

change distributions for both civilian and military rotorcraft. The shaded regions pertain to civilian 

rotorcraft, while the solid lines show boundaries for military design requirements and mishap data. 

The velocity profiles for military aircraft are much higher than civilian with the inclusion of MILSTD-1290 qualified rotorcraft. A velocity was chosen that would not extend beyond the civilian 

envelope. By also taking into account the achievable swing velocities of the LandIR, the vertical 

and horizontal velocities selected were 26 ft./sec and 35 ft./sec, respectively. A pitch attitude was 

chosen that would guarantee the test article would impact the surface at a nearly flat or slightly 

nose-up attitude. The accuracy of the pitch attitude for LandIR full-scale swing testing is ±1 

degree. A nose-up attitude of 2 degrees was selected. 

29
Mishap data has also revealed that nearly 75% of mishaps occur on soil terrain or water for both 

civilian and military rotorcraft. Compared to impact on a prepared surface like concrete, impact 

on soil would potentially decrease the vehicle vertical deceleration due to soil compliance, but 

would increase the longitudinal deceleration because of plowing and lesser rebound. Evaluation 

of the occupant and cargo restraints was predicated on introducing a substantial component of 

longitudinal deceleration (>10 g). A soil impact surface was chosen that was a combination of sand 

and clay that had been used previously for land-landing tests of the NASA Orion Crew Module 

(ref. 13). A soil bed 120-ft. long, 20-ft. wide, and 2-ft. tall was laid on top of the LandIR concrete. 

Figure 25. Comparison of civil and military survivable velocities 

RESULTS 

Pre and Post Test Photographs 

Photographs of the TRACT 1 airframe in the release position and just following impact are shown 

in Figure 26. All external and onboard high speed video cameras properly triggered and recorded 

data through the swing and impact phase. The actual impact conditions are listed in Table 7. A

photo time lapse from an external high-speed camera is shown in Figure 27 and Figure 28. The 

airframe impacted the soft soil with a 2.5° nose-up pitch attitude, causing the aft frames to impact 

initially. Next, the airframe rotated allowing the forward cabin to impact the soil. The duration of 

the point of impact to cockpit contact was approximately 0.020-sec. At 0.1-sec, the helicopter 

rebounded slightly, while maintaining a slight nose-down pitch and rolling towards the starboard 

side. A secondary impact occurred at 0.55-seconds, and the test article came to rest at 0.9-sec. The 

test article slide out was 8 ft. 

TRACT

30
Figure 26. Pre- and post-test photographs

Table 7. Impact conditions 

Design Actual

Vertical velocity 26 ft./sec 25-ft/sec

Horizontal velocity 35 ft./sec 33-ft/sec

Pitch 2° nose up 2.5° nose up

Roll 0° 0.5°

Yaw 0° Less than 1°

Figure 27. Impact sequence of TRACT I test- 0-0.200 seconds

(a) Pre-test photograph. (b) Post-test photograph. 

t 0 t 0 + 0.030 sec.

t 0+ 0.050 sec. t 0+ 0.100 sec.

Front cabin / Cockpit initial impact Close to cockpit max head displacement 

Mid cabin initial impact

31
Figure 28. Impact sequence of TRACT I test- 0.200-0.9 seconds 

Structural damage to the helicopter was observed in the shear panels located beneath the floor at 

discrete FS locations. Damage to these panels was more severe near the rear of the aircraft, which 

is the location of first ground contact. The rear shear panel deformations from FS 320 to FS 410 

are shown in Figure 29. Failure modes consisted of plastic deformation, crippling, and tearing. 

Minor deformation due to longitudinal shear was observed in the shear panels from FS 190 to FS 

286 (Figure 30). In addition, some tearing of the material was noted adjacent to rivet lines between 

the outer skin and fuselage frames. Wrinkling of the outer skin was also observed, as shown in 

Figure 31, of the skin underneath the cockpit. Minor deformation was evident on the frame sections 

above the floor aft of FS 286. No damage was evident on the cabin floor. 

t 0 + 0.450 sec.

t 0 + 0.550 sec. t 0 + 0.900 sec.

t 0 + 0.200 sec.

~Max rear rebound Rebound has flattened out

Rear second impact Airframe at rest

32
Figure 29. Photograph illustrating failure in rear cabin shear panels 

FS 382

FS 350

FS 320

FS 410

Tearing

Crippling

33
Figure 30. Photograph illustrating failure in mid-cabin shear panels 

Figure 31. Photograph illustrating skin wrinkling underneath cockpit 

34
Airframe Time History Results 

A schematic of the fuselage highlighting airframe FS locations is shown in Figure 32. All airframe 

acceleration traces are filtered using an SAE CFC60 filter (ref. 14). It is important to note that FS 

410 is located at the aft cabin/tail splice frame, FS 254 is located at mid-cabin, and FS 152 is 

located at the cockpit/forward cabin splice frame. The pilot and co-pilot responses are recorded on 

the floor supporting the seat rails. 

Figure 32. Schematic of helicopter showing instrumentation locations 

Given the nose-up pitch attitude at impact, it is expected that the onset of acceleration would occur 

first for FS 410, next for FS 254, and last from FS 152. The data shown in Figure 33 confirm this 

expectation. The left side acceleration responses range in peak magnitude between 25- and 45-g 

with durations of approximately 0.08 seconds. The right side responses range in peak magnitude 

between 22- and 55-g with a duration of approximately 0.08 seconds. The difference in responses 

between left and right are attributed to the seat m. At FS 152, there is a noticeable negative 

component of acceleration before 0.03 seconds as the test article pitches down. The behavior is 

more pronounced in the pilot and co-pilot responses. The co-pilot and pilot responses have very 

high magnitude oscillations after slam down at 0.04 seconds. There is a high different between the 

co-pilot and the pilot responses because of the locations of the accelerometer blocks on the 

thin-walled floor. 

FS410

FS254

FS152 Co-Pilot

Pilot

35
Figure 33. Vertical floor-level acceleration results for the left and right sides of the cabin 

and the cockpit 

Horizontal acceleration time history responses are shown in Figure 34, along with a fuselage 

schematic showing instrumentation locations. Three plots are shown for the left and right sides of 

the airframe, and for the cockpit. Each of the curves shows a similar response, with average 

magnitudes of approximately 10-g. The entire aircraft begins to longitudinally decelerate within 

the first 0.005 seconds. The magnitude of the peak horizontal acceleration is approximately 30-g; 

however, most traces are lower. 

The orientations and magnitudes of the resultant accelerations just before and after cockpit belly 

impact are illustrated in Figure 35. In Figure 35a, the orientation of the airframe deceleration is 

pointed to the aft and down. The pilot and co-pilot excursions are in the opposite direction, upward 

and forward. The magnitude of the negative component is approximately 10-g during nose over. 

In Figure 35b, after cockpit belly impact, the acceleration is directed up and slightly aft with a 

magnitude greater than 95-g. 

36
Figure 34. Horizontal floor-level acceleration responses 

Figure 35. Cockpit resultant acceleration 

Cabin mid-wall and floor vertical acceleration responses are shown in Figure 36, along with 

photographs indicating the locations of instrumentation. The mid-wall traces indicate two peaks, 

the first having a magnitude of approximately 20-g and the second having a magnitude of 25-g. 

The floor traces are from the two pair of forward facing seats. Whereas the mid-wall traces were 

37
very similar, the floor traces are opposite to one another. The forward seat exhibits a 15-g uniform 

acceleration response, which suddenly increases near the end of the pulse to a peak of 38-g. 

Conversely, the rear seat exhibits an initial peak of 34-g, which is reduced to a relatively uniform 

15-g response that decays near the end of the pulse. Both responses have durations of 

approximately 0.08 seconds. 

Figure 36. Cabin vertical acceleration responses at seat attachment locations 

In order to assess the spatial dependency of the accelerometer time histories, a metric proposed by 

Horta et al (ref. 14) uses singular value decomposition derived basis vectors, or impact shapes. In 

this approach, time histories from analysis or experiments are decomposed as: 

ݕሺݔǡ ݐሻ ൌ σ ߪ௜߶௜ሺݔሻ݃௜ሺݐሻ ௡

௜ୀଵ (1) 

In this form, the impact shape vector Ii sized m x 1 contains the spatial distribution information 

for m sensors, g(t) contains the time modulation information, V contains scalar values with shape 

participation factors, and n is the number of impact shapes to be included in the decomposition, 

often truncated based on allowable reconstruction error. The fractional contribution of the i

th

impact shape to the total response is proportional to δi, defined as: 

5

38
ߜ ௜ൌ ఙ೔

σ ఙ೗ ೙

೗సభ

 (2) 

When visualized, the impact shape vectors are also helpful in identifying responses that may be 

erroneous due to signal corruption, polarity, or sensor mount failure. The impact shapes are plotted 

by connecting each sensor location with a wireframe. The impact shape mesh and the first three 

impact shapes are plotted in Figure 37. The first two impact shapes are dominated by cockpit 

response. The first 10 impact shapes provide 80% of the overall contribution of the response. 

Figure 37. Impact shapes from test data 

Test Objective 1 Results for CH-46 Crew seat with MA-16 inertia reel and with PARS 

A sequence of images from the co-pilot and pilot cameras is shown in Figure 38 and Figure 39. As 

the rear of the cabin impacts, the PARS system immediately activates and 2-3 inches of webbing 

retracts within 0.010 seconds. During pitch-over, the pilot and co-pilot ATD excursions are 

forward and upwards. The feet are lifted off the floor pans, and their buttocks are no longer in 

contact with the seats. The restraints grab the ATDs and prevent significant excursion. At 0.040 

seconds, the cockpit belly impacts and the ATDs re-contact the seat cushions and foot pan. 

The seats vertical EAs begin to stroke at 0.080 seconds and the ATDs commence forward flail. 

Figure 40 shows the lumbar load filtered to SAE Class 600 and vertical pelvic acceleration filtered 

to SAE Class 1000. Peak lumbar load and vertical pelvic acceleration occur at the time of stroke. 

Post-test inspection of the seats indicated that the vertical and horizontal EAs stroked 

approximately 4 inches and 1.5 inches, respectively. 

The peak lumbar load is 1,500 lb. for the co-pilot and 2,200 lb. for the pilot, with the duration of 

loading about 0.060 seconds. While the timing of the EA stroke and motion of the pilot and co-pilot 

flail are qualitatively consistent, there could be slight differences in motion that would explain the 

forces. The PARS shoulder belts retracted about 3-4 inches at the point of aft cabin impact, 

restraining the co-pilot excursion more than the pilot during pitch-over. The altered positions 

100

150

200

250

300

350

400

-20

0

20

40

-40

-30

-20

x, [in]

y, [in]

z, [in]

11

13

15

16

17

18

19

20

21

22

23

24

25

29

27

30

100

150

200

250

300

350

400

-20

0

20

40

-40

-30

-20

x, [in]

y, [in]

z, [in]

100

150

200

250

300

350

400

-20

0

20

40

-40

-30

-20

x, [in]

y, [in]

z, [in]

100

150

200

250

300

350

400

-20

0

20

40

-40

-30

-20

x, [in]

y, [in]

z, [in]

Impact Shape Mesh (G=25%)

(G=18%) (G=8%)

39
between pilot and co-pilot when the cockpit belly impacted would have changed the distribution 

of loading through the lower leg and the pelvis. No tibial load cells were included in the 

instrumentation; therefore, those loads cannot be differentiated. The current FAA requirement for 

passenger seats certification is 1,500 lb. for a 50th percentile ATD. This criterion has been 

reconsidered by the DOD and a lumbar tolerance of 2,065 lb. was recommended (ref. 16). The 

injury tolerance for the pilot and co-pilot is exceeded based on the FAA requirements, but not the 

DOD recommendation. 

The pelvic acceleration waveforms track the lumbar load similarly. Magnitudes do not track 

linearly with lumbar load, with magnitude spikes above 70 g for the co-pilot and 100-g for the 

pilot. These abrupt peaks are caused by the ATD re-contacting the seat pan, and therefore not 

evident as the load path is damped into the lumbar. The relative velocities between the ATD and 

seat at the time of re-contact are different from the PARS retraction, causing different dynamic 

overshoot effects. These accelerations well exceed human acceleration tolerance limits for eyeballs 

down (+25 g) and would be categorized using Eiband curves as severe (ref. 17). The Dynamic 

Response Index in the spinal direction (DRIz) is 27.4 for the co-pilot and 30.2 for the pilot, 

indicating a >50% risk of AIS 2+ injury (ref. 18). These criteria are no longer recognized 

requirements, and have been superseded by lumbar criteria when evaluating crash loads 

specifically, but underscore the severe response of the pilot and co-pilot. 

The potential for significant neck injury due to axial tension and compression and flexion and 

extension bending are evaluated by plotting the upper neck axial forces (Fz) and moments (My). 

Maximum neck flexion occurs at approximately 0.120 seconds. The Neck Injury Criteria (Nij) is 

used to determine the combined effect of axial and bending loads (ref. 19). Kite graphs showing 

the acceptable limits for force and moment are shown in Figure 41. The time of maximum axial 

load and bending loads often do not coincide, so Nij is computed for multiple instances during 

impact. Nij is calculated for four select times for both pilot and co-pilot. For the pilot, all forces 

and moments are within acceptable limits, while there is one case for the co-pilot where the axial 

tension is marginal. 

40
Figure 38. Co-Pilot ATD response

Figure 39. Pilot ATD response

T-0.001 sec. T+0.045 sec. 

T+0.085 sec. T+0.123 sec. 

T-0.001 sec. T+0.045 sec. 

T+0.085 sec. T+0.123 sec. 

41
Figure 40. Pilot/Co-pilot ATD lumbar load and vertical pelvic acceleration 

Figure 41. Pilot/ Co-pilot Nij 

Test Objective 2 Results for Floor Mounted Passenger ATDs 

The sequences of images for the forward facing 50th percentile ATDs and 5th and 95th ATDs are 

shown in Figure 42 and Figure 43, respectively. The ATDs press into the seat cushions instantly 

after impact. Neck flexion and torso flail initiate at approximately 0.040 seconds after impact, and 

the maximum flail occurs at approximately 0.30 seconds. The ATDs rebound at 0.5 seconds, but 

flail a second time due to the test article rebounding and impacting again. 

Lap belt loads for all four ATDs were less than 350 lb. There was no evidence that energy 

absorption mechanisms, other than seat cushion compression, were actuated during the impact. 

Post-test inspection revealed no damage to the seat rail interface and insignificant permanent 

deformation within the seat frame. 

The lumbar loads are notably high for all four ATDs. Figure 44 shows the lumbar load for the 50th

percentile ATDs. Load waveforms are similar in shape and duration, which is expected given the 

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similarity in hardware. The peak lumbar load was 2,900 lb. for the 50th FAA Hybrid III and 

2,550 lb. for the 50th Hybrid II. The duration of the loading was approximately 0.060 seconds. 

These lumbar loads are well above the 1,500 lb. FAA limit, and would suggest a high likelihood 

of severe spinal injury. The peak lumbar load was approximately 1,600 lb. for the 5th Hybrid III 

and 2,500 lb. for the 95th Hybrid III. The lumbar load for the 5th female is an estimate from the 

waveform that appears to have saturated. This saturation is actually a scaling error identified 

post-test that limited the output range for that channel. The peak load of 1,600 lb. was determined 

through extrapolation. There is not a specific FAA requirement defined for the 5th and 95th ATDs. 

However, the DOD has recommended injury limits of 1,281 lb. for the 5th female and 2,534 lb. for 

the 95th male (ref. 14). The DOD recommendations for the 50th ATD are scaled up relative to the 

FAA requirement. Therefore, equivalent civilian requirements for the 5th and 95th would likely be 

lower. It is evident that lumbar loads for the 5th female and 95th male are near or above injury 

tolerance levels as well.

Figure 42. 50th ATDs response 

T-0.003 sec. T+0.058 sec. 

T+0.306 sec. T+0.582 sec. 

43
Figure 43. 5th and 95th ATD response

Figure 44. Forward facing ATDs, lumbar load 

T-0.001 sec. T+0.042 sec. 

T+0.286 sec. T+0.534 sec. 

44
Test Objective 3 Results for Standing ATD with Gunner's Belt and MARS 

The two mobile aircrew surrogates both released simultaneously from their temporary restraints 

and followed a symmetric trajectory during airframe deceleration to create relative velocity 

between the ATDs and the airframe. This continued until the ATDs were nearly in a “hands and 

knees” position on the floor of the airframe. Between steel cable release and the hands and knees 

position, the MARS locked and webbing payout was restricted. By the time the MARS ATD was 

in a hands and knees position, the MARS had arrested his vertical excursion toward the fuselage 

floor and he began to swing forward and up. At the same time, the gunner’s belt on the gunner’s 

belt ATD was still slack. The gunner’s belt ATD moved right past the hands and knees position to 

impact the cabin floor followed by a forward slide along the floor. Figure 45 contains a high speed 

video sequence from the outside of the aircraft showing key points during the crash sequence. 

Frame 1 is at the time the standing ATDs are released from their steel cable loops. At this point, 

the aft fuselage has contacted the soil, but the forward fuselage is still pitching down. In Frame 2, 

the head of the gunner’s belt ATD is visible inside the crew door. Frame 2 also represents the 

initiation of seat stroke for the cockpit occupants. Frame 3 clearly indicates a symmetric response 

between the gunner’s belt and MARS ATDs. The MARS reel has locked at a point between Frames

2 and 3. Also at the time of Frame 3, the cockpit crew seats continue stroking. Frame 4 represents 

the time of the end of seat stroke for the cockpit occupants. The mobile aircrew ATDs in Frame 4 

have symmetrically moved to the ‘hands and knees’ position. The MARS ATD has continued to 

move while pulling webbing off of the MARS reel by film spooling (approximately 8- 10 inches.). 

Frame 4 also indicates that the MARS ATD has started to slow relative to the airframe. Frame 5 

depicts a severe contact between the head of the gunner’s belt ATD and the floor of the airframe. 

No such contact has happened with the MARS ATD. Frame 6 shows the gunner’s belt ATD sliding 

forward on the floor and the MARS ATD pulling up away from the floor. Frames 7 and 8 depict 

the motion of the MARS ATD as he pivots about the MARS. 

45
Figure 45. Sequence of mobile aircrew motion during CH-46 airframe crash test. 

Figure 46a shows the resultant head acceleration time-histories for both ATDs. The head strike 

against the floor is clearly identifiable in the acceleration data. Both the magnitude and duration 

of this head strike are far in excess of human tolerance. With the MARS unit, the strike was 

prevented and head accelerations are substantially lower and sustained over a much longer 

duration. Head acceleration measurements indicate the head impact for the ATD with the gunner’s 

belt was a lethal impact. The calculated Head Injury Criterion (HIC) for the gunner’s belt ATD 

was 4,539. Correspondingly, the HIC for the MARS ATD was 248. For reference, HIC values of 

1,000 are used as an injury threshold for frontal automotive crash certification, FAA seat 

performance standards, and Navy qualification test requirements. Figure 46b contains resultant 

chest acceleration time-histories for the two ATD’s. Based on the typical 60-g critical chest 

acceleration limit, there is again an injurious situation with the gunner’s belt ATD and a noninjurious situation with the MARS ATD. Figure 47 shows measured neck forces and moments. 

Key in these graphs is the remarkable difference between the two ATD’s. For all neck forces and 

moments, the ATD with the MARS experienced less severe loading. 

46
Figure 46. Resultant Acceleration for Mobile Aircrew ATDs 

Figure 47. Neck Forces and Moments for Mobile Aircrew ATDs 

(a) Resultant Head Acceleration (b) Resultant Chest Acceleration

(a) Upper Neck Fz

(b) Upper Neck Mx (c) Upper Neck My

47
Test Objective 4 Results for Full-Field Photogrammetry 

The exterior photogrammetry data was captured and analyzed post-test. Data was computed such 

that a data point was acquired every 1.5 inches over the entire port side skin of the test article, with 

over 8,000 data points acquired. A sequence of photogrammetry data that were collected during 

the crash test is shown in Figure 48. The diagonal regions in which no data are shown are attributed 

to exterior cables, which obscure the dots. 

This displacement starts near the stub wing when this location comes in contact with the ground. 

As the point of impact of the test article moves forward during the impact, the computed out of 

plane motion also moves forward. These plots indicate flexing of the fuselage upon impact with 

the maximum out-of-plane (z) displacement of greater than 1 inch seen in the region of the middle 

round window above the swing beam between FS 286.0 and 320.0. Finally, the photogrammetric 

data show areas of permanent deformation around the stub wing locations and lower parts of the 

skin, near the floor. These locations match areas of damage seen in the fuselage post-test 

inspections taking the form of skin wrinkling and buckling. Figure 49 shows the discrete time 

histories of the out-of-plane points at WL +13. These points are interrogated further because the 

WL +13 points have the highest-on-average out-of-plane displacement. Figure 49 also shows 

vertical displacements taken from these same points at WL +44, which are the highest-on-average 

vertical values. These plots, taken together, suggest approximately 0.050 seconds after the impact 

shows the greatest fuselage deformation. 

Further evidence of this flexing is demonstrated in Figure 50. Vertical relative deformation is 

evaluated by interrogating relative position change between discrete points along the length of the 

cabin at the frame locations. Five locations were investigated further using this technique. The 

analysis calculates the maximum change in length of a created line endpoint positions, divided by 

the original undeformed line length, and reported as a percentage similar to strain calculations. 

The maximum relative deformation occurs in the front and mid cabin suggesting that the largest 

unsupported test article locations are those with the highest deformation. 

48
Figure 48. Measured out of plane displacement photogrammetry data 

Figure 49. Discrete point time-histories 

T+0.024 sec. T+0.036 sec. 

T+0.054 sec. T+0.076 sec. 

49
Figure 50. Relative vertical deformation photogrammetry data 

Test Objective 5 Results for Sidewall Mounted Troop Seats 

A sequence of images from the high speed camera aimed at the CACS side facing ATDs is shown 

in Figure 51. The ATDs began slumping into the seat bottoms at the point of impact. Vertical 

stroking of the seat EAs and head lateral rotation started just after 0.050 seconds. The peak head 

lateral bending occurs at 0.143 seconds, with higher rotation for the ES-2re head compared to the 

Hybrid III head. Both upper legs of the FAA Hybrid III shifted forward during the impact, twisting 

the body about the Z-axis. The feet of this ATD remained near their initial position during the 

impact. The Aerospace Hybrid-III’s entire right leg flailed forward, but the left upper leg was 

restrained from forward motion by the crotch strap of the 5-point restraint. This kept the torso of 

the ATD in its initial position, exhibiting very little Z-axis rotation. As the ATDs rebounded from 

maximum head lateral bending, the FAA Hybrid III returned to a side-facing orientation but the 

right upper leg of the Aerospace Hybrid III remained in the extended position. This difference in 

leg flail and body Z-axis rotation may be related to the restriction of hip rotation produced by 

interaction between the ATD femur and the fixed pelvis flesh of the FAA Hybrid III. 

Lumbar load and vertical pelvic acceleration are plotted in Figure 52. Examination of the midwall 

acceleration near each seat (Figure 36) indicates that the aft seat was loaded just prior to the 

forward seat. If the lumbar load data is overlaid with this time difference removed (a 0.010 second 

phase shift), it becomes apparent that the onset rate and overall response are very similar. The 

lumbar load peaked at approximately 1,000 lb. for both ATDs, and pelvic acceleration were limited 

to roughly 20 g. These loads are regarded as moderate in severity. The total stroke for each CACS 

seat was noticeably different. The seat with the ES-2re ATD stroked 4 inches and the seat with the 

Hybrid III ATD stroked 1.2 inches. The load drop exhibited in both cases is evidence that some 

energy was absorbed by both seats. The larger force drop in the forward seat correlates to the 

longer stroke observed. The difference in seat stroke may be related to factors such as the loading 

and unloading behavior of the ATDs, variability in the stroking force produced by each seat’s 

energy absorbers, or interaction with the standing ATD that had moved back against the forward 

seat occupant after release. Of note is that both seats limited spinal loads (as intended) even though 

the vertical component of the acceleration pulse applied to them at the sidewall interface was very 

50
different from the idealized triangular shaped pulse applied in dynamic tests to qualify seats of this 

type (Figure 53). 

The differences in stiffness for the ES-2re and Hybrid III necks are evident when comparing ATD 

upper neck forces and moments, as seen in Figure 54. Axial force and lateral bending moments 

were 50% lower for the ES-2re neck than the Hybrid III neck. The all of the loads measured at the 

ES-2re upper neck were well below the force and moment limits cited in the side-facing occupant 

neck injury criteria recently established by the FAA. These criteria include 405 lb. tension and 

compression, 185 lb. lateral shear, and 1,018 in-lb. lateral bending moment. 

Figure 51. Response of side facing ATDs on CACS 

T-0.001 sec. T+0.055 sec. 

T+0.143 sec. T+0.461 sec. 

51
Figure 52. Side facing ATDs on CACS, pelvic acceleration and lumbar load (HIII offset by 

0.010 seconds) 

Figure 53. Comparison of sidewall acceleration pulses 

52
Figure 54. Side facing ATDs on CACS, upper neck forces and moments 

The high definition camera aimed at the Hybrid II ATDs on the CH-46 troop bench shifted during 

impact, and was unable to capture video coverage of the ATD motion. Post-test photos (Figure 55)

showed that the seat fabric was torn off the aft frame, the front frame buckled near its center, and 

the right frame member detached from the aft frame. Because of the seat collapse, evaluation of 

lumbar load for injury is not appropriate, with more upper leg contact forces than pelvic. The ATD 

pelvises and upper legs were wedged within the frames, indicating a high likelihood of upper leg 

injury. Abdominal injury was also possible since the lap belt angle changed. 

53
Figure 55. Side Facing ATDs on troop bench

Test Objective 6 Results for Markerless Tracking of ATD 

A limited number of frames of ATD motion were captured using the markerless tracking system.

This result is partially due to its low acquisition rate of 30 frames per second, but primarily due to 

the foot of the standing ATD impacting and dislodging the sensor from its mount during the 

impact. This impact caused the sensor to break free of its mounted location and become settled on 

the floor of the airframe, pointed down toward a linkage. However, approximately 62 frames of 

color and depth images were captured during the free fall prior to impact. The time history in 

Figure 56 shows a summary of the data captured during the test. Each pair of images contains data 

from the color sensor, which is presented on the left, and a fringe plot reconstruction of the depth 

data, which is presented on the right. Since data were unavailable post-impact, the data presented 

occurs at four different states prior to impact. The depth data shows a graphic representation of 

ATD distance from the sensor, displayed in units of millimeters (mm). At pre-release, the majority 

of the ATD is colored green and blue, which represents 1,300 (blue) to 1,700 (green) mm position 

away from the sensor. The frame prior to impact, labeled pre-impact, shows the ATD is partly 

green, but primarily red in color. This data suggests that the ATD is between 1,700 and 2,200 mm 

away from the sensor, representing a half meter change in position away from the sensor from 

release to impact. 

54
Figure 56. Markerless tracking time history-color image (left) and depth fringe plot (right) 

The skeletal tracking algorithms did not associate with the depth data, and therefore, skeletal data 

was not acquired. Two potential reasons are presented to explain the lack of skeletal capture. The 

first reason was the shape of the standing ATD. The standing ATD contained a helmet, flight vest, 

and other various straps and articles which would cause confusion in the tracking algorithms. This 

theory was proven, when, during a preliminary test, this equipment was removed and the sensor 

was able to associate a skeletal wireframe with the standing ATD, as shown in yellow in Figure Pre-release Mid swing Pre-impact Post-impact

55
57. The second reason was a potential interference of the projected dot field due to the outdoor 

environment. Preliminary testing showed the infrared light generated from the sun had the 

potential to overlap the projected infrared dot field, essentially washing out the dots and obscuring 

the measurements. Large holes in the spatial data on the ATD are evident in Figure 56, presumably 

from this phenomenon. Unfortunately, the placement of the sensor could not be changed due to 

the high density of experiments on the test article. The position chosen behind the front bulkhead 

near the floor was the only acceptable location to locate the sensor for ATD tracking and to not 

interfere with other experiments. 

Figure 57. Preliminary ATD wireframe tracking 

Test Objective 7 Results for Non-Energy and Energy Absorbing Cargo Restraint 

Results for restraint forces in the standard webbing cargo restraint and the load-limiting cargo 

restraint are shown in Figure 58. The standard webbing cargo restraint developed a tensile load of 

nearly 5,000 lb. However, there was not enough kinetic energy in the cargo mass for the standard 

webbing to exceed the breakaway fitting load. The cargo mass was also decelerating in the vertical 

and longitudinal directions, which differs from typical sled test evaluations. The mass on the 

standard cargo restraint came to rest without severing the fitting, within 0.060 seconds. The loadlimiting cargo restraint exhibited load-limiting behavior about the design load of 3,000 lb. The 

tensile force plot revealed a loading and unloading behavior as the strap sections ripped and 

unfolded. The duration of the deceleration was extended to greater than 0.100 seconds. 

Skeletal Wireframe

56
Figure 58. Cargo Restraint Load 

Test Objective 8 Results for Three-Tiered Litter 

A sequence of images from the high speed camera aimed at the patient litter is shown in Figure 

59. During the swing phase, the unloading of the tie-down straps from freefall allowed the upper 

manikin to shift. The manikin’s head was partially suspended off the edge at impact. The manikin 

and ATD in the middle and lower litter did not move, since the tie-down straps remained in contact 

with the lower torso. At T+0.060 seconds, there was considerable deformation in the litter covers 

and longitudinal litter tubes. At T+0.080 seconds, the ceiling mounted strap that supports the aft 

inboard corner of the upper litter frame detached at the frame, and the litter stack up collapsed. 

The upper manikin’s head impacted the floor and rotated inward and forward with his legs still 

held with the tie-down straps. The middle ATD fell onto the lower manikin, and flailed freely 

forward after slackening of the tie-down straps. The lower manikin is impacted by the middle ATD 

and only shifted slightly forward. Severe injury is likely for all three ATDs. 

57
Figure 59. Patient Litter Sequence 

CONCLUDING REMARKS 

The TRACT 1 test was conducted in August 2013 at the LandIR facility. The impact test conditions 

(33-ft/s forward and 25-ft/s vertical velocity with a 2.5° nose-up pitch onto soft soil) were 

considered lower than typical DOD qualification levels, but severe enough to approach civilian 

requirements. The scenario tested has several parallels in field mishap data, and is believed to be 

a realistic and plausible condition that represents real world events. 

The TRACT 1 test was a highly successful collaboration between the FAA, DOD, and industry. 

The comprehensive data acquisition and photogrammetry were unprecedented in scope. Over 350 

channels of airframe and ATD data were collected with minimal loss of signal. External and 

onboard high speed and high definition cameras numbering more than 40 cameras provided for 

coverage not previously feasible in such a test. In addition to airframe structural responses, a total 

of 8 unique experiments were conducted as part of the crash test. These experiments included: 

1. Comparison of ATD responses in a CH-46 crew seat with MA-16 inertia reel versus a CH46 crew seat with PARS 

T-0.002 sec. T+0.060 sec.

T +0.080 sec T +1.000 sec.

58
2. Comparison of floor-mounted passenger seat and ATD responses 

3. Comparison of standing ATD with a traditional gunner’s belt versus a standing ATD with 

Aircrew Endurance Vest and MARS 

4. Full-field three-dimensional photogrammetry data collection 

5. Comparison of sidewall-mounted CACS troop seated Hybrid III ATD, seated Hybrid III 

ATD with ES-2re head and neck, and two seated Hybrid II ATDs in a standard CH-46 

tube and rag sidewall troop seat 

6. Evaluation of markerless tracking of ATD response 

7. Comparison of cargo experiment with non-energy-absorbing restrained cargo mass and 

energy absorbing restrained cargo mass 

8. Three-tiered litter mounted in the aft starboard region of the aircraft cabin. 

The horizontal decelerations were consistent throughout the aircraft at approximately 10-g with a 

duration of 80 msec. The vertical decelerations within the cabin varied from 20-50 g. The 2.5-

degree pitch up attitude caused the cockpit to accelerate downward just prior to belly contact. 

Vertical seat pan accelerations exceeded 100-g due to the weaker cockpit structure. 

ATDs were oriented in forward facing, side facing, and standing positions. The lumbar loads were 

within acceptable injury limits for the ATDs located in the pilot and co-pilot seats and two side 

facing CACS seats. The lumbar load for the co-pilot ATD with a PARS active restraint was lower 

than the lumbar load for the pilot ATD restrained using a standard MA-16 inertia reel. Lumbar 

loads exceeded injury limits for the four ATDs seated in forward facing passenger seats. While 

lumbar loads were low for the ATDs seated in the legacy troop bench, severe injury was likely due 

to tearing of the seat fabric and buckling of the frame. ATDs in the legacy patient litter would have 

likely sustained severe injury due to failure of the litter stanchions and collapse. Loads for the 

standing ATD held by the MARS system were within survivable limits, while fatal head blunt 

trauma occurred for the standing ATD held by the legacy gunner’s belt.

The stitched cargo restraint performed well and provided load limiting under longitudinal 

acceleration. Airframe skin vertical and lateral deformations during impact were computed with 

high-speed photogrammetry. This technique provides a more comprehensive method than 

employing strain gages or displacement transducers. Critical locations with high deformation 

could be readily identified over the entire fuselage skin. The markerless tracking experiment was 

able to provide position during the swing, but the impact of the ATD on the sensor prevented data 

acquisition during the impact. 

Airframe and experimental data for the TRACT 1 test will be compared to a second test of a CH46E containing similar experiments and conducted under similar test conditions, except that the 

airframe will be retrofitted with composite energy absorbing subfloor concepts. The TRACT 2 test 

will be conducted in late 2014. System-level tests like TRACT have the potential to enhance 

crashworthy system design, improve current regulations, and enable the next generation of aircraft 

airframe design. 

59
REFERENCES 

1. NASA Rotary Wing Project Overview, Briefing to NASA Langley Research Center, 

Susan Gorton, November 2012. 

2. Anonymous, Military Standard, MIL-STD-1290A (AV), Light Fixed- and Rotary-Wing 

Aircraft Crash Resistance, Department of Defense, Washington DC, 20301, September 

26, 1988.

3. Labun, L., “A Study of Rotary-Wing Crashes to Support New Crashworthiness Criteria,” 

AHS Forum 66, Phoenix, AZ, May 10-13, 2010. 

4. Kent, R., “Injury and Fatality Patterns in US Navy Rotary Wing Mishaps: A Descriptive 

Review of Class A and B Mishaps from 1985 to 2005,” AHS Forum 65, Grapevine, TX, 

May 27-29, 2009.

5. Couch, M. and Lindell, D., “Study on Rotorcraft Safety and Survivability,” AHS Forum 

66, Phoenix, AZ, May 10-13, 2010. 

6. Bolukbasi, A., Crocco J., Clarke, C., Fasanella, E., Jackson, K., Leary, P., Labun, L., 

Mapes, P., McEntire J., Pellettiere, J., Pilati, B., Rumph, F., Schuck, J., Schultz, M., 

Smith, M., and Vasquez, D., “Full Spectrum Crashworthiness Criteria for Rotorcraft,” 

RDECOM TR 12-D-12, December 2011. 

7. ADS-11B, Aeronautical Design Standard, Survivability Program, Rotary Wing, May 

1987.

8. Code of Federal Regulations - Title 14: Aeronautics and Space, Part 29: Airworthiness 

Standards: Transport Category Rotorcraft, 29.562 - Emergency landing dynamic 

conditions. 

9. Pellettiere, J., and Moorcroft, D., “Aircraft Seat Certification by Analysis from a 

Regulatory Perspective,” Proceedings of the American Helicopter Society 69th Annual 

Forum, Phoenix, AZ, May 21-23, 2013. 

10. DeWeese, R., Moorcroft, D., Abromowitz, A., and Pellettiere, J., “Civil Aircraft SideFacing Seat Research Summary,” DOT/FAA/AM-12/18, November 2012. 

11. Code of Federal Regulations - Title 14: Aeronautics and Space, Part 25: Airworthiness 

Standards: Transport Category Airplanes, 25.562 - Emergency landing dynamic 

conditions. 

12. Podob, R., Happ, M., and Sample, A.,“Mobile Aircrew Restraint System – MARS,” 

Proceedings of the Forty Third Annual SAFE Association Symposium, Salt Lake City, 

UT, October 24-26, 2005. 

13. Thomas, M., Chitty, D., Gildea, M., and T’Kindt, C, “Constitutive Soil Properties for 

Unwashed Sand and Kennedy Space Center,” NASA/CR-2008-215334, July 2008. 

14. Society of Automotive Engineers, Recommended Practice: Instrumentation for Impact 

Test – Part 1, Electronic Instrumentation, SAE J211/1, March 1995. 

15. Horta, L. G., Lyle, K. H., Lessard, W. B., “Evaluation of Singular Value Decomposition 

Approach for Impact Dynamic Data Correlation,” NASA TM 2003–212657, October 

2003.

60
16. Richards, M., Podob, R., “Development of an Advanced Energy Absorber,” SAFE 

Symposium Proceedings, 1997. 

17. Eiband, M. A., "Human Tolerance to Rapidly Applied Accelerations: A Summary of the 

Literature," NASA Memorandum 5–19–59E, June 1959. 

18. Brinkley, J. W. and Shaffer, J. T., "Dynamic Simulation Techniques for the Design of 

Escape Systems: Current Applications and Future Air Force Requirements," Symposium 

on Biodynamic Models and their Applications, AMRL–TR–71–29, Wright-Patterson Air 

Force Base, Dayton, OH, Aerospace Medical Research Laboratory, 1970. 

19. Eppinger, R., Sun, E., Kuppa, S., and Saul, R., “Supplement: Development of Improved 

Injury Criteria for the Assessment of Advanced Automotive Restraint Systems – II,” 

National Highway Traffic Safety Administration, March 2000. 

61
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Evaluation of the First Transport Rotorcraft Airframe Crash Testbed

(TRACT 1) Full-Scale Crash Test

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Annett, Martin S.; Littell, Justin D.; Jackson, Karen E.; Bark, Lindley W.;

DeWeesse, Rick L.; McEntire, B. Joseph

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NASA Langley Research Center

Hampton, VA 23681-2199

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National Aeronautics and Space Administration

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14. ABSTRACT

In 2012, the NASA Rotary Wing Crashworthiness Program initiated the Transport Rotorcraft Airframe Crash Testbed (TRACT) research program by obtaining 

two CH-46E medium lift helicopters from the U.S. Navy. The first crash test (TRACT 1) was performed at NASA Langley Research Center's Landing and 

Impact Research Facility (LandIR) in August 2013. Crash test conditions represented a severe but potentially survivable impact scenario. A total of 8 unique 

experiments were conducted to evaluate ATD responses, seat and restraint performance, cargo restraint effectiveness, patient litter behavior, and 

photogrammetric techniques. A combination of Hybrid II, Hybrid III, and ES-2 Anthropomorphic Test Devices (ATDs) were placed in forward and side 

facing seats, and in standing positions using mobile aircrew restraints. Occupant results were compared against injury criteria. Loads from ATDs in energy 

absorbing seats and restraints were within injury limits. Severe injury was likely for ATDs in forward facing passenger seats, legacy troop bench seats, and a 

three-tiered patient litter. In addition to occupant loading, the structural response of the airframe was assessed based on accelerometers located throughout the 

airframe and using three-dimensional photogrammetric techniques. Analysis of the photogrammetric data indicated regions of maximum deflection and 

permanent deformation.

15. SUBJECT TERMS

Energy absorption; Impact testing; Smulation

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