Silvia Seng and Charles Frankenberger

Naval Air Warfare Center Weapons Division, China Lake, California

Charles R. Ruggeri, Duane M. Revilock, J. Michael Pereira, and Kelly S. Carney

Glenn Research Center, Cleveland, Ohio

William C. Emmerling

FAA William J. Hughes Technical Center, Atlantic City International Airport, New Jersey

Dynamic Open-Rotor Composite Shield Impact

Test Report

NASA/TM—2015-218811

November 2015

DOT/FAA/TC–15/22
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Silvia Seng and Charles Frankenberger

Naval Air Warfare Center Weapons Division, China Lake, California

Charles R. Ruggeri, Duane M. Revilock, J. Michael Pereira, and Kelly S. Carney

Glenn Research Center, Cleveland, Ohio

William C. Emmerling

FAA William J. Hughes Technical Center, Atlantic City International Airport, New Jersey

Dynamic Open-Rotor Composite Shield Impact

Test Report

NASA/TM—2015-218811

November 2015

DOT/FAA/TC–15/22

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
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NASA/TM—2015-218811 1 

Dynamic Open-Rotor Composite Shield Impact Test Report 

Silvia Seng and Charles Frankenberger 

Naval Air Warfare Center Weapons Division 

China Lake, California 93555 

Charles R. Ruggeri, Duane M. Revilock, J. Michael Pereira, and Kelly S. Carney 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

William C. Emmerling 

FAA William J. Hughes Technical Center 

Atlantic City International Airport, New Jersey 08405 

Summary 

The Federal Aviation Administration (FAA) is working with the European Aviation Safety Agency to 

determine the certification base for proposed new engines that would not have a containment structure on 

large commercial aircraft. Equivalent safety to the current fleet is desired by the regulators, which means 

that loss of a single fan blade will not cause hazard to the aircraft. NASA Glenn and Naval Air Warfare 

Center (NAWC) China Lake collaborated with the FAA Aircraft Catastrophic Failure Prevention Program 

to design and test a shield that would protect the aircraft passengers and critical systems from a released 

blade that could impact the fuselage. This report documents the live-fire test from a full-scale rig at 

NAWC China Lake. NASA provided manpower and photogrammetry expertise to document the impact 

and damage to the shields. The test was successful: the blade was stopped from penetrating the shield, 

which validates the design analysis method and the parameters used in the analysis. Additional work is 

required to implement the shielding into the aircraft. 

1.0 Introduction 

The Federal Aviation Administration (FAA) is working with the European Aviation Safety Agency to 

determine the certification base for new engines being proposed for large commercial aircraft that would 

not have a containment structure. Regulators want the new systems to have safety equivalent to that of the 

current fleet, meaning that the loss of a single fan blade will not cause a hazard to the aircraft. One 

method to achieve this is to provide a shield that would protect passengers and the aircraft’s critical 

systems from a released blade that could impact the fuselage. To evaluate the feasibility of this concept, 

engineers from the NASA Glenn Research Center and the Naval Air Warfare Center (NAWC) China 

Lake collaborated with the FAA Aircraft Catastrophic Failure Prevention Program to design and test a 

composite shield that could provide the required protection. The test was conducted in February 2014 by 

the Naval Air Warfare Center Weapons Division (NAWCWD) Weapons Survivability Lab (WSL) on 

behalf of the FAA Aircraft Catastrophic Failure Prevention Program. The purpose of the test was to 

evaluate the effectiveness of composite panels in preventing penetration of the aircraft fuselage by an 

open-rotor aircraft engine blade. Glenn engineers designed and analyzed the composite shields and the 

test configuration and provided photogrammetry expertise, equipment, and analysis support. The 

LS–DYNA explicit finite element code (Livermore Software Technology Corp., Livermore, CA) was 

used for the shield impact analysis. NAWCWD WSL conducted the test and provided the test hardware. 

Based on an analysis performed by NASA (Ref. 1), two shield thicknesses were tested, 20 and 24 ply. 

The blades were predicted to penetrate the 20-ply composite material, but not the 24-ply composite 

material. Two blades, similar in size and construction to that expected for open-rotor engines on a 
NASA/TM—2015-218811 2 

midsize commercial aircraft, were mounted on a spin fixture and run at a representative velocity. Using a 

linear shape charge, the two blade sections were released simultaneously so that they would impact the 

composite shields. The test was designed to duplicate the worst-case scenario; that is, the blades were 

released at a high-power-climb rotor speed, with both blade sections impacting tip first and with the blade 

aligned with the trajectory of the center of gravity (CG) of each blade. The test was conducted 

successfully, and the results were consistent with the NASA analysis predictions: The thinner shield was 

penetrated by the blade, and the thicker shield prevented penetration. This report documents the live-fire 

test from the full-scale rig at NAWC China Lake, describes the damage to the shields, and presents the 

instrumentation results. Test results are reported in the units they were measured in. 

1.1 Background 

In the 1980s, aircraft engine manufacturers began to investigate new engine design concepts to 

improve fuel efficiency because of rising fuel costs resulting from the oil embargo in 1973, the revolution 

in Iran, and the Iran-Iraq war. One of the concepts was an open-rotor unducted turbofan. Open-rotor 

engines may have the potential to reduce the specific fuel consumption by as much as 25 percent or more. 

Two main open-rotor concepts were developed with the goal of achieving an engine that would be able to 

provide comparable aircraft speeds to a ducted turbofan at much greater efficiencies (Ref. 2). Pratt & 

Whitney, Hamilton Standard, and Allison teamed up to offer the 578–DX, a geared propfan with 

counterrotating blades. The General Electric (GE) unducted fan (UDF) also used two sets of 

counterrotating blades (Fig. 1). Rather than a traditional stator-rotor turbine, the GE UDF engine was 

designed with counterrotating turbines as well, each one providing power to one set of UDF blades. 

Two major technical challenges—significant noise generation and exposure to uncontained propeller 

debris—had not been solved in earlier designs. These challenges and reduced fuel prices resulted in 

decreased interest to develop the open-rotor design concept in the 1980s. 

Current increased fuel costs and greenhouse gas environmental concerns have reinvigorated interest 

in open-rotor and UDF engine designs. The FAA Office of Environment and Energy, NASA, and engine 

manufacturers have been evaluating new engine designs that utilize two unshrouded counterrotating rows 

of blades. Significant improvement in engine noise reduction has been achieved by advances in computer 

modeling for blade design and use of composite materials for open-rotor blades. 

Use of advanced composites in fuselage structures may address airplane threats from uncontained 

blade failures. This report documents results of testing to validate the effectiveness of composite panels 

for shielding the fuselage from debris. 

The open-rotor engine design concept does not fall under existing regulations for either turboprop or 

turbofan engine installation. The fan blades are mounted external to the engine flow path, and there are 

many more blades than there are on a turboprop. The renewed interest in the development of open-rotor 
NASA/TM—2015-218811 3 

engines and unducted turbofan engine designs has led the FAA to investigate the feasibility of fuselage 

shielding for protection in the event of a blade release. The goal of the FAA is for open-rotor engines to 

have an equivalent level of safety as current ducted fan engines where loss of a single blade is not a 

hazardous event. 

The FAA initiated an open-rotor shield research program that consisted of the following sequence of 

events: 

1. NASA performed penetration tests on high-damage-tolerance braided composite materials (Ref. 3). 

2. NAWCWD conducted generic open-rotor-engine aircraft studies to show the benefits of a composite 

shielding material (App. A). 

3. NASA conducted detailed shield penetration analyses to investigate the feasibility of composite 

shielding (Ref. 1). 

4. NAWCWD fabricated and tested a composite shield design (this report). 

5. NAWCWD conducted a cross engine debris trade study (App. B). 

Glenn engineers conducted analyses and small-scale testing to predict the thickness needed for a 

composite material to stop a released fan blade. The aircraft geometry used for the analyses was based on 

a medium-range, twin-engine, midsize (150-passenger) narrow-body commercial aircraft with a high 

wing and a wing-mounted open-rotor engine. The open-rotor engine specifications were provided by 

NAWCWD and the FAA. The configuration incorporated open-rotor, wing-mounted engines as shown in 

Figure 2. This generic configuration does not address the cross-engine debris certification requirement 

discussed in Appendix B. 

The diameter of an open-rotor engine suitable for this size of an aircraft was to be between 13.1 and 

13.8 ft based upon Reference 2. For the test, a 162-in.- (13.5-ft-) diameter rotor was selected based on 

available test assets and hardware. The horizontal distance between the centerline of the engine and the 

fuselage centerline was 188 in., with a vertical distance of 85 in. 

The analysis was conducted in two parts. In the first part a trajectory analysis was conducted. The 

aircraft and engine geometries were modeled to determine what release angles would result in the worstcase impacts to the fuselage. The release angles providing a velocity vector perpendicular to the fuselage 

skin and having the highest impact velocity were found to be a release angle of 30 (blade rotating 

counterclockwise) or 180 from the top (blade rotating clockwise), as shown in Figure 2. The full details 

of the analysis are described in Reference 1. 

The results of the first part of the analysis were used to determine the position and velocity of the 

blade just before impact. These parameters were used as inputs to a finite element model for the second 

part of the analysis. The computational time was reduced by positioning the blade at the predicted 

position just before impact, rather than beginning at the engine release point. The finite element analysis 

indicated that a 24-layer (approximately 0.52-inch-thick) T700S/PR520 triaxially braided composite was 

sufficient to stop the blade from penetrating the shield. 
NASA/TM—2015-218811 4 

1.2 Test Objective 

The following tasks were to be performed in this test: 

1. Release representative blade fragments into two braided composite panels and collect data that can 

be used to validate the LS–DYNA analysis of composite panels to protect the fuselage from a fan 

blade-off event. 

2. Document the released blade trajectory to obtain data that can be used to validate the NASA 

trajectory study and answer questions raised by industry on previous large angular trajectory 

experience. 

2.0 Test Article Description 

The test involved two main components, the composite fuselage panel and the simulated open-rotor blade. 

2.1 Composite Fuselage Panel 

Analysis of shielding materials to protect a fuselage from a single blade was conducted by the NASA 

Glenn Research Center Impact Dynamics Group. The composite fuselage panel built up for this test series 

was based on the NASA design analysis (Ref. 1). The panel did not contain any stiffeners or stringers. 

The design concept for the shielding is a floating panel that will not be subject to the flexure of the 

primary aircraft structure. By isolating the panel from the fuselage structure, the panel is not exposed to 

those additional stresses and strains and can be made of a lighter weight material than a structural 

shielding panel. 

The composite test panel was a curved 4- by 8-ft section with a radius of curvature of 6.5 ft. Four 

composite panels were fabricated: two composed of 20 layers of braided carbon fiber prepreg, each 

resulting in a thickness of approximately 0.45 in., and two composed of 24 layers, each resulting in a 

thickness of approximately 0.56 in. The material was a prepreg made with a triaxial carbon fiber (T700S) 

braid from A&P Technology, Inc. (see Fig. 3). The prepreg resin was Cytec MTM45–1 resin (Cytec 

Industries, Inc.). The resin was chosen because it provides a high-strength, damage-resistant, structural 

epoxy matrix that can be processed in an autoclave without expensive tooling and was similar to the 

PR520 used in the analysis. The specifications for the material are shown in Table I. 

The composite panels were made by the NAWCWD composite materials shop on tooling designed and 

built by WSL range engineers and the WSL shop. The composite panel represents the composite barrier 

only, with no stringers. The panel was supported in a frame that matched the NASA-modeled boundary 

conditions as closely as possible. The panel mounting structure, described in Section 5.2, was designed to 

simply support the panel in the three translational directions, allowing rotation along the panel edges. 

TABLE I.—FEDERAL AVIATION 

ADMINISTRATION OPEN-ROTOR PROJECT 

COMPOSITE MATERIAL SPECIFICATIONS 

Preform 

A&P product code QISO–H–52.8–A 

Width 52.8 in. 

Fiber areal weight 536 gsm 

Fiber orientation 0  60

Directional content 33.3 percent in each orientation 

Fiber type T700SC–12K50C 

Prepreg 

Resin MTM45–1 

Resin content 36 wt% 

Width 52.8 in. 

Process method Dual film 
NASA/TM—2015-218811 5 

2.2 Rotor Blade 

The fan blades of an open-rotor turbofan are large, adjustable-pitch, unshrouded blades. The design 

typically features two rows of counter-rotating blades. The unshrouded blades provide greater efficiency 

than traditional turbofans. Noise issues that hindered the development of these engines in the 1980s have 

been reduced with the help of advances in computational fluid dynamics (CFD) programs. 

The blades used for this testing (see Figs. 4 and 5) were similar in design and materials to those that 

could be used for open-rotor engines. A modern propeller was used as the surrogate with the outer portion 

acting as the open-rotor fan blade-off simulation. The blades have a polyurethane foam core, sandwiched 

between carbon fiber spars, with composite reinforced skins. A thin strip of metal protects the leading 

edge from small, “nominal,” foreign object damage, such as small debris from unimproved runways 

(see Figs. 4 and 5). 
NASA/TM—2015-218811 6 

3.0 Test Approach 

A single test event was conducted. Two blades were mounted to the rotor hub and spin rig on the 

WSL spin fixture, shown in Figure 6. The spin fixture is a T406 engine and gearbox that are capable of 

providing the 1108 rpm speed required for this test. The two blades were explosively released at 8 past 

vertical. Two 4- by 8-ft composite panels were located approximately 124 in. from either side of the 

centerline of the spin fixture, so that they were centered on the predicted blade release trajectories. The 

fuselage structure was not represented. 

Two of the six blades were mounted in the spin rig. Linear shape charges were placed 41.25 in. from the 

blade tip of each blade. The shape charges were simultaneously initiated to release the blades at the same time. 

3.1 Blade Separation Investigations 

Prior to the start of testing, a pretest investigation was conducted to validate the ability to cut the 

blade with minimal damage to surrounding blade structure and imparting minimal force on the resulting 

blade section. An additional objective for this testing was to assess the potential for damage to the spin 

fixture and/or damage to the released blade section that may influence test results. 

A single blade was fixed to an overhead mounting structure (Fig. 7). The blade was not allowed to 

rotate. A weight was suspended from the end of the blade. The suspended weight was meant to encourage 

separation, but was much smaller than the centripetal forces on the blade at 1108 rpm (200 lb for the 

pretest investigation and over 26 000 lb estimated for the blade in motion). 

A witness board was placed adjacent to the blade to represent the spin fixture. The damage to the 

witness board was used to evaluate the damage from the detonation of the linear shape charge. A 

600 grain/ft linear shape charge was placed along the chord of one side of the rotor blade, closest to the 

 
NASA/TM—2015-218811 7 

witness board. If the shape charge had effectively severed the blade, without significant damage that could 

influence test results, the test would have been deemed successful. However, the shape charge did not fully 

sever the blade. The spar adjacent to the linear shape charge was severed, but the opposite spar was intact. 

Because the shape charge did not completely sever the blade, the test was repeated with the shape 

charge cord around the full circumference of the blade (Fig. 8). The second test was performed on the 

same blade, several inches above the location of the first test. The blade was completely severed by the 

linear shape charge cord around the full circumference of the blade. The spars appeared to have visibly 

delaminated some on either side of the linear shape charge. A large piece of the blade trailing edge 

(aft of the carbon spars) separated from the remaining blade section; however, this appears to be a result 

of damage from the first test. 

Test instrumentation was limited to one high-speed video camera, a safety video, and a firing line. 

The shape charge resulted in a significant fire ball, but it did not impart significant force on the blade. 

Figure 9 shows the blade root section rotating as a result of the blast, but does not propel the blade 

vertically. Likewise, the outer span of the blade merely falls away from the cut. Review of the severed 

blade section reveals a relatively clean cut with composite delamination reaching approximately 1 in. 

away from the cut line (Fig. 10). 
NASA/TM—2015-218811 8 

Additional pretesting was done after the T406 spin fixture and test rig were installed onto the test pad. 

Pretest checkouts included engine runs as well as testing of the linear shape charge firing system. 

Detonators were placed on the fan blades but not the linear shape charge. The pretest was conducted just 

as the final full-up test would be, utilizing the same procedure. The flash from the detonators functioning 

is visible in the high-speed video. The detonator flash provided verification for the timing of the firing 

signal without damaging the blades and demonstrated the integrity of the firing chain. 

The detonators were inserted into a small aluminum cube to secure the detonator and the end of the 

linear shape during the test. During one of the pretest spin-ups without the detonators installed, the 

detonator block broke free from the blade and impacted one of the test panels. Although there was no 

visible damage from the detonator holding block, the panel was replaced. Detonator blocks were then 

bolted onto the blade to ensure they remained in place. 

3.2 Test Pad Setup 

The T406 engine was mounted to the WSL engine test stand with the open-rotor fan blade-off test rig 

attached as shown in Figure 11. 

The test pad setup was based on the worst-case trajectories defined by the NASA analysis (Ref. 1). To 

minimize the height of the fuselage panel fixture, the centerlines of the fuselage and the engines were 

placed at the same height. This orientation results in a 24 rotation of the coordinate system, centered on 

the centerline of the fuselage (shown in Fig. 12). Therefore, the release angle of a blade at the 180

orientation (blade oriented straight down) would become 204. The 30 release angle would become 6

(30 – 24 = 6). 
NASA/TM—2015-218811 9 

The release angle for the test was chosen to be 8. This was based on the recommendations from 

NASA to provide the worst-case impact angle for the blades simultaneously released from the hub. The 

8 release angle was predicted to result in the tip’s first impact with the blade to be aligned with the 

velocity vector of the blade CG. The composite test panels were placed such that the centerline of the 

fuselage panels and engine centerline were at the same height. The horizontal distance of the panels from 

the engine was set to ensure that the distance from the release point CG to the impact on the composite 

test panels was the same as the distance used by the NASA analysis. The final placement of the panels 

relative to the engine is illustrated in Figure 13. 

The panels were held in a test fixture that replicated the boundary conditions of the NASA analysis as 

shown in Figures 14 and 15. The panels were fixed in all three translation axes and simply supported on 

half-round bars that allowed the edges to rotate about the support frame interface (Fig. 15). The panels were 

restrained along all four edges utilizing 3/8-in. bolts, spaced approximately every 4 in. around the frame, and 

tightened down with sufficient force to restrain the panel in the translation axis. A rubber gasket was 

inserted between the restraint bars and the composite panels to minimize stress concentrations in the panels. 

The linear shape charge was installed to cut the blade 41.25 in. in from the blade tip. The spin fixture 

was operated at 1108 rpm to provide the target blade velocity of 527 ft/s. This provided the target blade 

weight of 15.1 lb and release blade CG of 26 in. from the tip (Table II). 

 
NASA/TM—2015-218811 10 

Table II.—TARGET BLADE CHARACTERISTICS 

Release blade length, in. ..................................... 41.25 

Release blade weight, lb .......................................15.1 

Release center of gravity diameter, in. .................. 109 

Rotor speed, rpm ..................................................1108 

Release angle, deg .................................................... 8 

Blade velocity, ft/s ................................................ 527 
NASA/TM—2015-218811 11 

3.3 Video Recording 

NASA personnel provided cameras for digital image correlation (DIC) and personnel to set up, 

operate, and provide the analysis for the DIC. The imaging system was the primary data source to 

evaluate the displacement of the test panels. These data were required to validate the NASA analysis of 

the composite panels. 

The six cameras provided by NASA were operated in pairs. Two cameras were placed behind each test 

panel to record the blade impact. The third camera pair was placed forward of the test fixture, looking aft, to 

capture the release of the blade and impact on the composite panel. The cameras were located behind steel 

blast shields to protect them from debris from the impact of the blade into the panels. Lighting was provided 

to illuminate the fuselage panels to ensure the maximum frame rate and desired shutter speeds could be 

achieved. The cameras were placed in fixed positions and calibrated to provide the DIC. The DIC software 

was used to process video images from which the displacement of the panel was calculated. 

Two cameras were provided by the WSL to capture overall views of the test event. One camera was 

placed in front of the test fixture looking aft, a second high-speed camera was placed aft looking forward. 

The approximate camera placement is shown in Figures 16 to 18. Additional standard-speed (24 frames/s) 

video cameras were placed on the test pad at the discretion of the firing officer to monitor safety views of 

the test pad and to monitor the spin fixture engine. 
NASA/TM—2015-218811 12 
NASA/TM—2015-218811 13 

3.4 Instrumentation 

High-speed video was used to measure the fan blade time-of-flight. Sensors installed on the spin 

fixture monitored the health of the engine and measured the spin rig velocity (in rpm). Additional data 

were provided by 12 accelerometers on the panels and along the restraint bars. The measurements made 

during the test are listed in Table III. 

The accelerometers were installed in the same locations and orientations on each panel support 

structure; their locations on the 20- and 24-ply panels are shown in Figures 19 and 20, respectively. For 

each panel, three accelerometers were placed in the lower left corner of the frame (viewed from the 

concave side of the panel): (1) P20–X1 and P24–X1 for the 20- and 24-ply panels, respectively, are 

aligned with the fuselage (“x-axis”); (2) P20–Y1 and P24–Y1 are aligned in the tangential (“y-axis”) 

orientation; and (3) P20–Z1 and P24–Z1 are aligned in the radial (“z-axis”) orientation. Two 

accelerometers were placed in the lower right corner of the panels (viewed from the concave side of the 

panel): (1) P20–Y2 and P24–Y2 aligned in the tangential orientation and (2) P20–Z2 and P24–Z2 aligned 

in the radial orientation. Finally, one accelerometer was placed in the upper left corner: P20–Y3 and 

P24–Y3 aligned in a near-tangential orientation, as shown in Figure 21. The accelerometers were used to 

collect backup data to compare with the NASA analysis data. The accelerometer data are presented in 

Appendix C. 

TABLE III.—TEST MEASUREMENTS 

Parameter Units Data rate 

Gas generator speed Percent of maximum limit 

Power turbine speed Percent of maximum limit 20 Hz 

Propeller speed Revolutions per minute (rpm) 

Torque Pounds-foot (lb-ft) 

Main gearbox temperature Degrees Celsius (C) 20 Hz 

Engine fuel flow Gallons per minute (gpm) 

Gearbox oil pressure Pounds per square inch, gauge (psig) 20 Hz 

Oil temperature 

Bearing temperature Degrees Fahrenheit (F) 20 Hz 

Bearing temperature 

Zero time signal Volts (V) 100 kHz 

Acceleration, 20-ply panel 

 P20–X1 

 P20–Y1 

 P20–Z1 

 P20–Y2 

 P20–Z2 

 P20–Y3 

Acceleration of gravity (g) 100 kHz 

Acceleration, 24-ply panel 

 P24–X1 

 P24–Y1 

 P24–Z1 

 P24–Y2 

 P24–Z2 

 P24–Y3 

Acceleration of gravity (g) 100 kHz 
NASA/TM—2015-218811 14 
NASA/TM—2015-218811 15 

3.5 Photogrammetry 

Three pairs of high-speed digital video cameras were used to quantitatively measure the full-field 

deflection of the back side of each test panel as well as the kinematics of one of the released fan blades. 

The commercial DIC software ARAMIS (GOM mbH, Germany) was used to compute the threedimensional panel deflections and in-plane strains based on images obtained from a calibrated pair of 

cameras focused on the back side of each panel. The panels were painted with 0.625-in.-diameter black 

dots, on 1-in. centers, on a white background. For the 20-ply panel a pair of Photron model SA5 cameras 

(Photron USA, San Diego, CA) were mounted on a tripod behind the panel in such a way that the whole 

panel was visible. These cameras had a resolution of 832 pixels along the length of the panel and 600 

pixels across the width, a frame rate of 15000 frames/s and an aperture time of 1/59000 s. For the 24-ply 

panel, a pair of Photron model SA1.1 cameras was used, in a similar configuration. The camera resolution 

along the length of the panel was 704 pixels and 448 pixels across the width. The frame rate was 18000 

frames/s and the aperture time was 1/49000 s. The system used on the 24-ply panel is shown in Figure 22. 

To measure the kinematics of the blade impacting the 20-ply panel, the commercial software 

PONTOS (GOM mbH, Germany) was used in conjunction with a pair of Phantom V7.3 cameras (Vision 

Research Inc.). The camera pair was mounted forward and to the right of the engine axis viewing the area 

from the blade release to the panel impact. These cameras had a frame rate of 12500 frames/s, a 

horizontal resolution of 768 pixels, a vertical resolution of 504 pixels, and an exposure time of 20 s. 

Figure 23 shows the camera pair location relative to the test hardware and Figure 24 shows a 

representative frame from the video after blade release and prior to impact. The white dots on the blade 

and the test article were individually tracked in order to obtain blade linear and angular velocity during 

the test. However, because of the large amount of light generated from the explosive charges used to 

release the fan blades and the resulting overexposure of the images, only a portion of the free flight of the 

blade segment could be tracked. 
NASA/TM—2015-218811 16 
NASA/TM—2015-218811 17 

3.6 Blade Release 

A linear shape charge was placed at the separation point of the blade. The shape charge was affixed to 

the fan blade by an adhesive and covered with a thin fiberglass strip, as described in Section 3.7.2. The 

fiberglass strip was used to ensure that the shape charge remained in the correct position even when the 

blades are at maximum velocity, while minimally impacting airflow. 

3.6.1 Blade Release Firing System 

Based on the analysis performed by NASA Glenn, the ideal release angle for the blades was predicted 

to be between 0 and 13 past vertical blade position. The WSL has developed a blade release firing 

system to provide precise release of the blades. This blade release firing system is designed to release the 

blades within 1 of the ideal release angle. 

The firing system for this test event had to provide a high-voltage current pulse to four explodingwire detonators, which initiated the linear shape charges that severed the blades. The firing pulse was 

timed to release the blades at the designated angle. This firing system had to meet the safety requirements 

defined in the WSL Standard Operating Procedures for Explosive Bridgewire Operations (EBW) 

requirements. A graphical representation of the firing system is shown in Figure 25. 

The blade position was detected by using a Banner Engineering Corp. model PD49 retroreflective 

laser sensor. This provided a blade “zero time” pulse to a timing box. A variable potentiometer in the 

timing box allowed manual adjustment of the delay from blade “zero” to a targeting camera strobe pulse 

and to a fire pulse logic circuit. The targeting camera was aligned with the desired firing position. 
NASA/TM—2015-218811 18 

The timing pulse was set by adjusting the potentiometer so that the target blade was correctly aligned 

within the camera view. The logic circuit would not pass the firing pulse until the facility sequencer 

provided the firing pulse. A targeting laser provided a visual aid for fine adjustment. When all conditions 

for a test were met, the facility time sequencer was started. When the time sequencer countdown reached 

zero, the fire pulse was allowed to propagate to the test pad electronics. 

Two electronic boxes sat on the pad. The high-voltage box provided 3000 to 3500 V. The highvoltage trigger box provided the electronic switch to control the high-voltage power supply. To reduce the 

vulnerability of the high-voltage trigger box electrical noise, optical isolators were used. Also, to prevent 

the system from firing if de-energized, uninterruptible power supplies were used for each element of the 

firing system. Additionally, a Safe/Arm relay was installed to interrupt the fire pulse at the isolation 

transformer until sequencer time zero, which will only allow the switch to function during the first 

revolution. Monitoring circuits were also installed to detect any problems in the test environment. 

The high voltage traveled through slip rings mounted on the propeller shaft to the exploding-wire 

detonators on the propellers. Multiple tests were conducted to test the firing system prior to the full-up 

blade release testing. The firing system functioned nominally during this test. 

3.6.2 Linear Shape Charge Installation 

Placement of the shape charge is indicated in Figure 26. Exploding-wire detonators were placed on 

either end of each linear shape charge. The two detonators provided redundancy, allowing the full length 

of the explosive to fire if any single wire had broken or if there had been a discontinuity in the linear 

shape charge. 
NASA/TM—2015-218811 19 

The linear shape charge has a foam adhesive strip to provide the proper standoff distance to the 

material to be cut. To attach the shape charge more securely to the blade, a layer of fiberglass was placed 

over the shape charge. Additionally, microballoon filler was used to provide support, and a second layer 

of fiberglass was applied (Figs. 26 and 27). 

4.0 Pretest Prediction(s) 

Pretest predictions were based on the analysis conducted by NASA Glenn Research Center, 

documented in Reference 1. Based on the analysis, the fan blade was not expected to penetrate the 24-ply 

(0.52-in.) composite panel, but was expected to fully penetrate the 20-ply composite panel. 

5.0 Test Results 

Testing was conducted on February 20, 2014. All facility systems functioned nominally, and the 

blades were released as intended, at 8 past vertical (Fig. 28). The test conditions are noted in Table IV. 

The blades separated as intended and struck the composite test panels. 

Figure 28 shows a composite photograph of the release of the blades just before shape-charge 

detonation, just after the shape charges were fired, and at first impact of the blades against the composite 

panels. The upper blade impacted the 24-ply panel and resulted in cracks, but no penetration. The lower 

blade impacted the 20-ply panel and, as predicted, penetrated the panel. The test results showed good 

correlation to the NASA analysis. 

Data collected from the high-speed cameras were analyzed by NASA. The results of the DIC analysis 

are presented in Section 5.3, Photogrammetry. 

Additional analysis was conducted to estimate the test impact condition. Blade-release velocity was 

analyzed by reviewing high-speed video and calculating velocity from engine revolutions per minute. 

Geometry of the test and timing from the high-speed video provided a calculated blade fragment velocity 

of 534 ft/s. A second calculation was conducted based on propeller speed (revolutions per minute) and an 

estimated blade fragment CG radius of 54.5 in. This calculation resulted in a blade velocity of 533 ft/s. 
NASA/TM—2015-218811 20 

TABLE IV.—TEST CONDITIONS 

Test condition Desired 

(24-ply/20-ply panel) 

Actual 

(24-ply/20-ply panel) 

Blade release position, deg 188/8 188/8 

Blade velocity, rpm 1108 1120 

Blade release velocity, ft/s 527 533 

Blade weight, lb 15.1 

Distance of blade impact (forward 

of plane of rotation), in. 5 12.5/11.5 

Blade kinetic energy, ft-lb 68 657.41 67 188.32 

20-ply panel thickness, in. 0.452 

24-ply panel thickness, in. 0.560 

Blade fragment trajectory was also estimated by reviewing high-speed video. The blade segments 

were released at 8 and 188 from vertical. From video analysis, the blade fragments traveled on a 

trajectory 4 less than the 8 and 188 perpendicular paths, resulting in the blade fragments impacting the 

panels above and below, respectively, the intended impact point. The blade trajectory obliquity was also 

calculated from video images. Figure 29 shows the impact obliquity for the upper-release blade is 

estimated to be 34. Also shown is the obliquity of 14 that is predicted for an 11 release angle. 
NASA/TM—2015-218811 21 

5.1 Lower (20-Ply) Composite Panel Test Results 

The release angle of the lower blade was 188. The blade released cleanly from the hub and rotated 

about the CG of the released fragment. The trajectory of the blade was approximately 4 above horizontal. 

The blade impacted tip first, with the length of the blade along the velocity vector. The blade caused a large 

tear in the panel, approximately 45 in. vertically and 12 in. horizontally, and penetrated completely (Figs. 30 

to 33). Two additional horizontal cracks, visible in Figure 30, caused delamination but did not penetrate the 

full thickness of the composite. After passing through the panel, the blade impacted the mount frame and 

deflected upwards from the test panel mounting fixture. Upon recovery of the blade on the test pad, the tip 

was missing and the residual length of the blade was only 36 in. (Fig. 34). 

The back side of the panel had a matching crack vertically but had additional vertical and horizontal 

cracks along the bottom of the panel and up both sides of the panel that did not penetrate the full thickness 

of the panel (see Figs. 31 and 33). 

The accelerometer block on the lower left side of the panel mounting fixture was dislodged and the 

P20–Y1 wire was cut during the impact and penetration of the blade through the material. Data from 

P20–Z2 spikes to 1500 g approximately 9 ms after apparent impact, possibly indicating that debris 

impacted the accelerometer. The sustained high-g data (1500 g) for P20–Y2 is indicative of possible 

damage to the sensor, which was installed in the same mounting block as P20–Z2. 

Ultrasound nondestructive inspection (NDI) of the 20-ply composite panel, shown in Figure 35, 

indicates a large area of delamination. The signal transmission scale is shown with the figure. No 

transmission at all (indicating delamination) is shown as white. The red dots throughout the green 

background are noise and are not indicative of damage. The small white holes along the border of the scan 

are the bolt holes for mounting the panel to the test fixture. The large white area on the scan indicates a large 

area of delamination, which corresponds to the visible cracks in the composite panel. The small patches of 

signal transmission within the large delaminated area may indicate partial damage or “kissing disbands”; 

however, this could not be determined from the ultrasound NDI. A thin line just below the bottom 

attachment of the panel to the fixture may indicate damage due to the mounting structure. A thin white line 

protruding from the lower left area of the delaminated area was caused by a thin piece of material protruding 

from the crack, which blocked transmission of the signal, not a delamination. 

The residual velocity of the blade after penetration of the panel was not determined because of the 

limited view of the back side of the composite panel. The tip of the blade was missing. The trailing edge 

was also damaged, but it is unknown at what point this damage occurred: upon impact with the panel or 

subsequent impacts with the lighting frame or the test pad. Material from the blade was wedged into the 

crack in the panel and was removed prior to NDI. 
NASA/TM—2015-218811 22 
NASA/TM—2015-218811 23 
NASA/TM—2015-218811 24 
NASA/TM—2015-218811 25 

 

 
NASA/TM—2015-218811 26 

5.2 Upper (24-Ply) Composite Panel Test Results 

The release angle of the upper blade was 8. The blade released cleanly from the hub and rotated 

about the CG of the released fragment. The trajectory of the blade was approximately 4 below 

horizontal. The blade impacted tip first, with the length of the blade aligned with the velocity vector. The 

blade caused tears on both the front and back of the panel, but the tears were not aligned with one another 

and did not penetrate the panel completely. The damage to the composite panel is shown in Figures 36 to 

39. Tears on the front side are visible in Figure 36, and tears on the back side are indicated by the arrows 

in Figure 37. 
NASA/TM—2015-218811 27 
NASA/TM—2015-218811 28 

The blade deflected upwards after impact and landed on the test pad. The tip and leading edge of the 

blade were crushed on impact, and the remaining tip and leading edge had no residual strength (Fig. 40). 

The residual length of the blade was 41 in. and weighed 11.2 lb. 

Figure 41 shows the results of the NDI scan of the 24-ply panel. No transmission at all (indicating 

delamination) is shown as white. The red dots throughout the green background are noise and are not 

indicative of damage. The small white holes along the border of the scan are the bolt holes for mounting 

the panel to the test fixture. The large white area on the scan indicates a large area of delamination, which 

corresponds to the visible cracks in the composite panel. 

 
NASA/TM—2015-218811 29 

5.3 Results: Photogrammetry 

Calibrated pairs of high-speed cameras and commercial DIC software were used to obtain 

quantitative measurements of blade kinematics and also panel displacements and strains. 

5.3.1 Blade Kinematics 

Photogrammetric measurements of the blade kinematics yielded the blade velocity, which impacted 

the 20-ply panel while in free flight. Figure 42 shows the resultant velocity as a function of time for the 

point estimated to be the blade CG (blue line) and a marker point near the blade tip (black line). There is 

some noise in the results, but the average resultant velocities over the measured time were 536 ft/s for the 

CG and 695 ft/s for the point near the tip. 

5.3.2 Panel Displacements and Strains 

Photogrammetric measurements of the back side of the 20- and 24-ply panels yielded measurements 

of the full-field displacements in three directions, and the resulting in-plane strains. For each test panel a 

local cylindrical coordinate system was established in which the origin was at the center of a cylinder that 

would be generated by extending the panel through 360. Positive values of radial displacement are 

outward from the origin. With the given camera orientations, positive axial (y-direction) displacement is 

in the engine forward direction for the 20-ply panel and in the engine aft direction for the 24-ply panel. 

The results indicated that there was measurable deflection in the frame to which the panels were clamped. 

It was possible to measure the panel deflection relative to the frame by subtracting the average frame 

movement. 
NASA/TM—2015-218811 30 

5.3.2.1 20-Ply Panel 

The 20-ply panel was penetrated by the blade. At the point of maximum deflection prior to failure the 

radial displacement at the point of impact was 1.175 in. The radial displacements for the 20-ply panel are 

shown in Figure 43. Figure 43(a) shows a contour plot of the radial displacement just prior to failure, with 

positive values being inward in the general direction of blade travel. Figure 43(b) shows the deflection 

along the segment line shown in part (a) for a number of consecutive frames up to the point of failure. 
NASA/TM—2015-218811 31 

Maximum principal strains for the 20-ply panel are shown in Figure 44. Figure 44(a) shows the 

distribution of maximum principal strain in the panel just prior to failure. At the point shown in the figure, 

the value was just over 1 percent. 

The time history of maximum principal strain of the point indicated in Figure 44(a) is shown as a 

function of time in part (b) up to the point of failure. In the figure, blade impact begins at 0.000 s. Prior to 

this time there is a strain response likely related to the blast generated from the shape charge, which 

travels faster than the blade. After penetration occurred, tracking the points was no longer possible. 
NASA/TM—2015-218811 32 

5.3.2.2 24-Ply Panel 

The 24-ply panel was not penetrated by the blade, making it possible to measure displacements and 

strains at their maximum values. 

5.3.2.2.1 Radial Displacements 

The blade did not penetrate the 24-play panel, but cracks were generated both on the front and back 

sides of the panel. Two main cracks were generated on the back side as shown in Figure 39. Radial 

displacements for the 24-ply panel are shown in Figures 45 and 46. Figure 45(a) shows a contour plot of 

the radial displacement at the time of maximum panel deflection. This is the total displacement, including 

any contribution from the motion of the test fixture. A number of points are indicated in the figure. The 

points labeled “0” and “1” span the first crack that formed on the back side of the panel shown in 

Figure 39 between 2 and 2.5 ms after impact. A time history of the radial displacement of three of the 

indicated points is shown in Figure 45(b). Note that during the time of impact the displacement at point 2 

is radially inward as would be expected, but the displacements at points 0 and 1 are outward indicating a 

highly dynamic process. The dashed vertical line in Figure 45(b) indicates the time at which the crack that 

formed between points 0 and 1 occurred. Figure 45(c) shows the radial displacement along the line 

labeled “Section 1” shown in Figure 45(a) for a number of consecutive frames up to the point of 

maximum deflection. Note that there is significant radial displacement at both ends of the segment line, 

which are near the fixture frame. This indicates that there is some motion in the frame itself. Although the 

motion of the frame is likely not a rigid-body motion, an attempt to measure the displacement of the panel 

relative to the frame was made by averaging the displacement of four points located at the corners of the 

measurable area and subtracting that motion from the overall deflection (Fig. 46). This measure of 

displacement relative to the frame is shown in Figure 46(a). Relative to the average of the four points, the 

maximum radial displacement is just over 2.6 in. This relative measure of radial displacement is shown as 

a function of time in Figure 46(b) for the seven points shown in part (a). There is a significant difference 

in the relative displacement shown in Figure 46(b) compared with the total displacement of the points 

shown in Figure 45(b). 

5.3.2.2.2 Transverse Displacements 

Displacements in the transverse (axial, y-) direction for the 24-ply panel are shown in Figures 47 and 

48. Figure 47(a) shows the y-displacement (upward in the figure) in the panel at the time that the first 

crack initiated. The y-direction is in the engine aft direction. Two points labeled “0” and “1” that span the 

crack are shown in the figure. The y-displacement of these two points is shown in Figure 47(b) as a 

function of time. The dashed line in the figure indicates the time of initiation of this crack. The figure 

shows that initially the two points labeled “0” and “1” are moving together. At the point of crack 

initiation, indicated by the dashed line, the displacement of point 0 becomes greater than point 1 in the 

positive y-direction, indicating that the two points are moving closer together and that the crack that 

formed at this location is the result of transverse compressive strain. At this time this region of the panel 

is displacing in the positive (outward) radial direction. 

Figure 48(a) shows the y-displacement of the panel near the time at which the second crack initiated 

near the panel boundary. This crack formed between the points labeled “7” and “8.” The figure shows that 

at this time there is generally an inward displacement of the panel near the upper and lower edges of the 

figure, possibly due to deformation of the frame fixture. The y-displacement of these two points is shown 

in Figure 48(b) as a function of time. In the figure, first impact occurs at 0.0000 s. The dashed line in the 

figure indicates the time of the initiation of the crack at this location. The figure shows that up until 

approximately 3 ms after impact, the y-displacement of the two points was essentially the same. After 

this, the displacement of point 8 was greater than that of point 7, indicating that the distance between the 

two points decreased. Right before failure, the relative displacement of the two points changed, and the 

distance between the two points began to increase, indicating that a tensile strain transverse to the crack 

location existed at the time of failure. 
NASA/TM—2015-218811 33 
NASA/TM—2015-218811 34 
NASA/TM—2015-218811 35 
NASA/TM—2015-218811 36 

 
NASA/TM—2015-218811 37 

5.3.2.2.3 Strains 

Strains in the 24-ply panel are shown in Figures 49 to 51. The strain in the y-direction (transverse to 

the cracks) at the time of the first crack initiation in the 24-ply panel is shown in Figure 49(a). In the 

region of the formation of the first crack, between the points labeled “0” and “1,” the strain in the 

y-direction is compressive. The y-direction strain at these points is plotted as a function of time in 

Figure 49(b). Because of the limited resolution of the cameras, it was not possible to measure the strain at 

the exact location of crack initiation. 

Figures 49(a) and (b) show that at the time of the first crack initiation this region was in a 

compressive strain state. 
NASA/TM—2015-218811 38 

Figure 50(a) shows a contour plot of the strain in the y-direction (transverse to the panel cracks) at the 

time the second crack initiated. The second crack initiated at the panel boundary, which was outside of 

the computable area, so it was not possible to compute the strain at the initiation site. However, as seen in 

the figure, the transverse strain is generally positive in this region, indicating that the crack initiated as a 

result of tensile transverse strain. Figure 50(b) shows the strain in the y-direction for the two points 

indicated in part (a), as a function of time. 
NASA/TM—2015-218811 39 

The maximum principal strain at the point of maximum deformation (indicated by point 2 in Figure 46(a)) 

is shown in Figure 51. No fracture occurred on the back side of this panel at this location. Unlike the 20-ply 

panel, there is no significant strain generated from the blast from the shape charge prior to impact. This is 

likely due to the differences in panel orientation. As shown in the figure, at the point of maximum deformation 

the panel undergoes a maximum strain of approximately 1.5 percent without fracture. 

5.4 Summary of Posttest Analysis 

The DIC results provide useful data for developing and validating computational models that simulate 

the impact of the blade on the shielding specimens and help to explain some of the mechanisms of failure. 

For the 20-ply panel deflections and strains up to the point of failure are provided. More information can 

be obtained from results from the 24-ply panel test. At the point of maximum deflection, the 24-ply panel 

sustained a maximum in-plane strain of approximately 1.5 percent without failure, whereas for the 20-ply 

panel the maximum in-plane strain was approximately 1 percent at the time of penetration. It was not 

possible with the DIC system to measure out of plane strains. The results indicate a highly dynamic event, 

with wave generation and fixture movement. The maximum radial displacement in the 24-ply panel is 

approximately 2.6 in. relative to the average displacement of four points at the corners of the specimen. 

The overall maximum radial displacement, including frame movement is approximately 4.3 in. 

Displacements of points spanning the two cracks that formed on the back side of the specimen indicate 

that the crack that initiated at the fixture frame was a result of transverse tension strains, while the crack 

that developed in the middle of the panel initiated in a region that was displacing radially outward and 

was a result of transverse compression strains resulting from the dynamic wave propagation in the panel. 

6.0 Conclusions 

The testing described in this report validated the results of the NASA LS–DYNA shielding design 

analysis. The results of the testing correlate very well with the NASA analysis results that predicted a 

penetration threshold between 20 and 24 plies of composite material for the open-rotor composite shield 

impact test. In addition, the test provided data on global deformation of the composite panels and local 

stress concentration leading to crack initiation. The blade release in the test occurred nearly as planned 

except that a NASA review of the blade trajectory determined that the blades impacted further forward 

than predicted because of the aerodynamics of the spinning blade not being taken into account. 
NASA/TM—2015-218811 40 

Since the intent of this test was to validate the NASA analysis, the test utilized the boundary 

conditions specified in the analysis. This test demonstrated that a relatively low-weight solution may be 

possible for open-rotor fuselage shielding and also validated tools that can be used to evaluate potential 

designs. The test also demonstrated the importance of considering local high-strain effects during impact 

that can lead to crack initiation away from the impact site. 

Numerical methods previously developed and the experimental methods described in this report provide 

tools that can be used for development and validation of future open-rotor fuselage shielding designs. 

7.0 Recommendations 

The results of this testing indicate that the NASA analysis methodology is a good tool to evaluate the 

use of barriers to protect the fuselage from uncontained blades. Additional investigation should be 

conducted, as design concepts for open-rotor aircraft emerge, to investigate the feasibility of protecting 

the fuselage from blades or blade fragments from these large rotors. 

 
NASA/TM—2015-218811 41 

Appendix A.—Open-Rotor Protective Shield Study 

The Naval Air Warfare Center Weapons Division (NAWCWD) was tasked under the Federal 

Aviation Administration’s (FAA’s) Catastrophic Failure Prevention Program to conduct an open-rotor 

protective shield study using the FAA Penetration Equations developed for the Uncontained Engine 

Debris Damage Assessment Model. The objective of this effort was to define the fuselage skin thickness 

required to stop impact from open-rotor blades and blade fragments. The tasking required the definition of 

a generic high-wing aircraft configuration with an open-rotor propulsion system (Fig. 52). Several 

iterations of the study were conducted with the initial effort defining a 2- by 8-ft shield of uniform 

thickness. Follow-on efforts defined optimized shield configurations and explored different shield 

materials. Fuselage structure was not considered in this analysis. 

Two generic blade configurations were used in this analysis (Fig. 53). An aluminum blade was used 

for reference purposes and a composite blade was defined to better represent current open-rotor 

configurations. The open-rotor system was sized based on input from industry (Table V). The blade 

parameters are provided in Table VI. The second rotor blade was roughly a 20-percent-scaled version of 

the first blade. Spacing between rotors was defined at 15 in. from rotor 1 centerline to rotor 2 centerline. 

The rotor-to-rotor distance for modern open-rotor engines was not publicly available when this work was 

initiated, so a distance similar to the 1980s unducted fan was used in this analysis. More recent 

information indicates that modern open-rotor engines have a rotor-to-rotor centerline distance 2 to 3 times 

that used in this analysis. This will increase shield area and weight beyond those published in this report. 

Variables in the study included blade material, fragment size, fragment orientation, and fuselage 

material. Blade and fuselage materials investigated were aluminum, a general carbon fiber/epoxy 

composite, and a braided carbon/high-strength epoxy fiber composite. Analysis was conducted to define 

the shielding weight required to stop the blade fragments. Results are provided for several shield 

configurations: a 2- by 8-ft uniform-thickness shield, an optimized shield for each individual rotor, and an 

optimized shield for both rotors. Optimized shield dimensions are based on fore-to-aft trajectory spreads 

of 3. A summary of the initial uniform shield study is presented in Table VII. This table shows the skin 

thickness required to stop a full blade fragment at a 90 (edge on) impact orientation for the various blade 

and fuselage materials. The weights listed are for an individual side of the aircraft so total aircraft shield 

weight would be twice the presented value. 

A follow-on study was conducted to assess an optimized shield weight for the generic composite 

fuselage shield configurations. For this study, the shield is defined by the material thickness beyond 

nominal (0.1 in.) needed to stop the blade fragment. Blade impact conditions (trajectory, impact obliquity, 

and fuselage curvature) play a significant role in this design. As impact obliquity angles increase, the line 

of sight thickness through the material increases so less material is needed to stop a given blade fragment. 

This concept also takes advantage of the close proximity of rotors 1 and 2, using only the thicker panel 

segment in the overlap area. An example of the optimized shield is provided in Figure 54. Study results 

are presented in Table VIII. The study investigated a generic composite fuselage with a composite blade 

fragment at a blade fore and aft trajectory limit of 3. Parameters included blade impact orientations of 

0 (flat), 60, and 90 (edge on); 1/3 blade fragment; and full-blade fragment for rotors 1 and 2. 

A final iteration of the shield study was conducted with a new high-strength, fracture-tolerant 

braided-composite fuselage material (Table IX). The study parameters were the same as those used in the 

generic composite effort. 

A shield of uniform thickness provides adequate protection at the cost of excess weight. An optimized 

shield significantly reduces excess weight and provides the same level of protection against the defined 

blade fragment. For the case of a full composite blade at 90 impact orientation, the generic composite 

shield of uniform thickness weighs 195.7 lb. By optimizing the shield thickness, the weight is 

significantly reduced to 65.8 lb, a 129.9-lb weight savings. Using higher strength materials provides 

added benefit. The braided-composite-optimized shield weighs 42.2 lb compared with the generic 

composite 65.8 lb, an additional 23.6-lb weight savings. 
NASA/TM—2015-218811 42 

TABLE V.—OPEN-ROTOR FIRST ROTOR 

OVERALL CHARACTERISTICS 

Diameter, ft ............................................... 13.42 

Blade length, ft ............................................ 3.44 

Hub diameter, ft .......................................... 6.54 

Rotor speed, rpm ........................................ 1020 

TABLE VI.—ROTOR BLADE GENERIC CONFIGURATIONS 

 Aluminum blade Composite blade 

Rotor 1 Rotor 2 Rotor 1 Rotor 2 

Full blade 

Length, ft 3.44 2.75 3.44 2.75 

Width, ft 1.0 0.8 1.0 0.8 

Average thickness, in. 0.8 0.8 2.0 2.0 

Average skin thickness, in. --- --- 0.3 0.3 

Weight, lb 39.6 25.3 17.0 10.8 

Centroid radius, ft 4.93 4.64 4.93 4.64 

Velocity, ft/s 527 497 527 497 

1/3 Blade 

Length, ft 1.1 0.88 1.1 0.88 

Width, ft 1.0 0.8 1.0 0.8 

Average thickness, in. 0.3 0.3 0.3 0.3 

Average skin thickness, in. --- --- 0.15 0.15 

Weight, lb 4.68 3.04 2.7 1.7 

Centroid radius, ft 6.71 6.16 6.71 6.16 

Velocity, ft/s 647 594 647 594 
NASA/TM—2015-218811 43 

TABLE VII.—UNIFORM-THICKNESS 2- by 8-FT SHIELD 

THICKNESSES AND WEIGHTS 

[Rotor 1, full blade, 90° impact orientation.] 

Blade material Fuselage material 

Aluminum General composite Braided composite 

Blade 1 

Aluminum 1.414 in. 

326.3 lb 

1.483 in. 

194.7 lb 

1.145 in. 

150.3 lb 

Composite 0.886 in. 

204.4 lb 

0.929 in. 

122.0 lb 

0.717 in. 

94.2 lb 

Blade 2 

Aluminum 1.182 in. 

272.7 lb 

1.239 in. 

162.7 lb 

0.957 in. 

125.7 lb 

Composite 0.731 in. 

168.7 lb 

0.561 in. 

73.7 lb 

0.433 in. 

56.9 lb 

 
NASA/TM—2015-218811 44 

TABLE VIII.—OPTIMIZED SHIELD GENERIC COMPOSITE STUDY RESULTS 

[Blade fore and aft trajectory limit of 3.] 

Rotor Fragment 

impacts 

on 

fuselage 

Baseline fuselage results Shielded fuselage optimized results 

Fuselage 

weight, 

lb 

Maximum 

residual 

velocity 

Fragment 

penetration, 

percent 

Fuselage 

weight, 

lb 

Optimized 

shield 

weight, 

lb 

Maximum 

required 

thickness, 

in. 

Shielded area 

in.2 ft2

(a) Full generic composite blade, composite fuselage 

Fragment impact 0

Combined 

rotors 

104 51.9% 21.3 0.476 2632.2 18.3 

Rotor 1 52 28.3 360.7 57.7% 42.5 14.2 0.476 1462.3 10.2 

Rotor 2 56 30.5 247.1 42.9% 37.5 7.0 0.359 1169.9 8.1 

Fragment impact 60

Combined 

rotors 

104 59.6% 33.0 0.593 3022.1 21.0 

Rotor 1 52 28.3 421.6 65.4% 50.3 22.0 0.593 1657.3 11.5 

Rotor 2 56 30.5 319.6 50.0% 41.5 11.0 0.424 1364.8 9.5 

Fragment impact 90

Combined 

rotors 

104 69.2% 65.8 0.929 3509.6 24.4 

Rotor 1 52 28.3 506.6 76.9% 74.0 45.7 0.929 1949.8 13.5 

Rotor 2 56 30.5 446.1 57.1% 50.4 19.9 0.561 1559.8 10.8 

(b) 1/3 generic composite blade, composite fuselage 

Fragment impact 0

Combined 

rotors 

101 41.6% 8.1 0.337 2047.2 14.2 

Rotor 1 52 28.3 312.9 46.2% 33.9 5.6 0.337 1169.9 8.1 

Rotor 2 52 28.3 245.4 34.6% 30.8 2.5 0.276 877.4 6.1 

Fragment impact 60

Combined 

rotors 

101 47.5% 13.3 0.389 2339.7 16.2 

Rotor 1 52 28.3 412.7 50.0% 36.9 8.6 0.389 1267.3 8.8 

Rotor 2 52 28.3 342.3 42.3% 33.0 4.6 0.318 1072.4 7.4 

Fragment impact 90

Combined 

rotors 

101 55.4% 23.4 0.482 2729.7 19.0 

Rotor 1 52 28.3 579.8 57.7% 42.7 14.4 0.482 1462.3 10.2 

Rotor 2 52 28.3 502.0 50.0% 37.3 8.9 0.394 1267.3 8.8 

 
NASA/TM—2015-218811 45 

TABLE IX.—OPTIMIZED SHIELD BRAIDED COMPOSITE STUDY RESULTS 

[Limit trajectory to ±3.] 

Rotor Fragment 

impacts 

on 

fuselage 

Baseline fuselage results Shielded fuselage optimized results 

Fuselage 

weight, 

lb 

Maximum 

residual 

velocity 

Fragment 

penetration,

percent 

Fuselage 

weight, 

lb 

Optimized 

shield 

weight, 

lb 

Maximum 

required 

thickness, 

in. 

Shielded area 

in.2 ft2

(a) Full generic composite blade, braided composite fuselage 

Fragment impact 0

Combined 

rotors 

104 44.2% 10.2 0.368 2242.2 15.6 

Rotor 1 52 28.3 333.5 50.0% 35.8 7.5 0.368 1267.3 8.8 

Rotor 2 56 30.5 206.1 35.7% 33.2 2.7 0.277 974.9 6.8 

Fragment impact 60

Combined 

rotors 

104 51.9% 18.3 0.458 2632.2 18.3 

Rotor 1 52 28.3 402.8 57.7% 41.3 13.0 0.458 1462.3 10.2 

Rotor 2 56 30.5 286.9 42.9% 35.7 5.2 0.327 1169.9 8.1 

Fragment impact 90

Combined 

rotors 

104 62.5% 42.2 0.717 3168.3 22.0 

Rotor 1 52 28.3 498.1 71.2% 58.8 30.5 0.717 1803.5 12.5 

Rotor 2 56 30.5 423.5 50.0% 42.0 11.5 0.433 1364.8 9.5 

(b) 1/3 generic composite blade, braided composite fuselage 

Fragment impact 0

Combined 

rotors 

101 24.8% 2.0 0.26 1218.6 8.5 

Rotor 1 52 28.3 248.5 34.6% 30.1 1.8 0.260 877.4 6.1 

Rotor 2 52 28.3 122.6 15.4% 28.5 0.2 0.213 390.0 2.7 

Fragment impact 60

Combined 

rotors 

101 35.6% 4.9 0.301 1754.8 12.2 

Rotor 1 52 28.3 359.1 42.3% 32.0 3.7 0.301 1072.4 7.4 

Rotor 2 52 28.3 255.8 26.9% 29.5 1.2 0.246 682.4 4.7 

Fragment impact 90

Combined 

rotors 

101 47.5% 11.6 0.372 2339.7 16.2 

Rotor 1 52 28.3 537.4 50.0% 35.9 7.6 0.372 1267.3 8.8 

Rotor 2 52 28.3 438.9 42.3% 32.2 3.9 0.304 1072.4 7.4 

 
NASA/TM—2015-218811 46 

Appendix B.—Cross-Engine Debris Study 

The Naval Air Warfare Center Weapons Division (NAWCWD) was tasked under the Federal 

Aviation Administration’s (FAA’s) Catastrophic Failure Prevention Program to conduct an open-rotor 

cross-engine debris study. The study was to be conducted using a generic low-wing midsize commercial 

aircraft configuration (Fig. 55); however, the results could be applied to a high-wing configuration 

currently used in many turbopropeller airplane configurations. The objective was to determine the design 

space (rotor vertical position) available for an open-rotor configuration to meet a 1/40 (9 of 

circumferential exposure) minimization requirement for cross-engine debris. 

The analysis was based upon using the inherent shielding provided by the fuselage to minimize the 

exposure of damage to the opposite engine following a blade release. The fuselage diameter is similar to a 

typical narrow-body turbofan-powered transport airplane such as the Boeing 737 or Airbus A320. The 

open-rotor engine diameters were selected based upon current conceptual engine designs. Two blade 

diameters were assessed. The baseline rotor system was 13.5 ft, and a second, longer blade configuration 

(16.5-ft diameter) was also assessed for comparison purposes. The two blade lengths and rotor parameters 

are defined in Table X. In the analysis, the vertical position of the rotor centerline was adjusted to the 

point where the blade fragment first intersected the second rotor. The rotor centerline height above 

fuselage centerline and circumferential release point were recorded. The rotor was then raised until the 

circumferential release point equaled the original release point plus 9. Again the rotor centerline above 

fuselage centerline was recorded. This analysis was conducted for a 1/3-blade fragment and a full-blade 

fragment. The critical parameters for the analysis were the center of gravity (CG) locations for the 1/3- 

and full-blade fragments, fuselage height, and rotor centerline height. 

TABLE X.—ROTOR DIMENSIONS USED IN 

CROSS-ENGINE DEBRIS STUDY 

Dimension Baseline blade Larger blade 

Diameter, ft 13.5 16.5 

Diameter, in. 162 198 

Tip radius, in. 81 99 

Hub diameter, in. 79.50 79.50 

1/3 blade length, in. 13.8 19.8 

1/3 blade center of gravity (CG) 

radius, in. 

74.1 89.1 

Blade length, in. 41.25 59.25 

Full blade CG radius, in. 55.0 79.0 

Weight, lb 15.1 21.7 

 
NASA/TM—2015-218811 47 

For the 13.5-ft-diameter system 1/3-blade analysis, the blade fragment initially impacts the second 

rotor when the rotor is raised 2 in. above fuselage centerline at a 359 release angle (Fig. 56). 

The maximum vertical rise that still meets the 1/40 requirement is a 32.1-in. rotor height above 

fuselage centerline (Fig. 57(a)). 

Analysis was conducted for the 13.5-ft-diameter rotor with a full-blade fragment. For the full blade, 

the initial impact to the second rotor occurs when the rotor is raised 11.5 in. above fuselage centerline. To 

meet the 1/40 requirement, maximum rotor system vertical rise could be up to 41 in. above fuselage 

centerline (Fig. 57(b)). 
NASA/TM—2015-218811 48 

A second analysis was conducted with a 16.5-ft-diameter open-rotor system. With the larger rotor 

diameter (increased blade length) the design trade space is reduced to 18 and 23 in. from fuselage 

centerline for the 1/3-blade fragment and full-blade fragment, respectively (Fig. 58). 

The objective of this analysis was to determine the design space (rotor vertical position) available for 

a low-wing open-rotor configuration to meet the 1/40 (9 of circumferential exposure) minimization 

requirement for cross-engine debris. Two rotor diameters were assessed—13.5 and 16.5 ft—and both 

systems had sufficient design space to meet the 1/40 cross-engine debris minimization requirement. As 

would be expected, the smaller diameter system had a larger trade space that met the requirement. A 

summary of results is provided in Table XI. 

TABLE XI.—RESULTS SUMMARY, ENGINE 

CENTERLINE VERTICAL TRADE SPACE AVAILABLE 

Rotor diameter, ft. 1/3 blade, in. Full blade, in. 

13.5 32.1 41 

16.5 18 23 

 
NASA/TM—2015-218811 49 

Appendix C.—Test Stand Vibration Data 

Six accelerometers were placed on each of the test stands supporting the test articles as previously 

described in Section 3.4. The intent of these accelerometers was to provide boundary condition 

information for future finite element analysis. In this appendix, digital image correlation (DIC) 

information is also provided to identify motion of the shields at the corners of the stand in addition to the 

accelerometer data. 

The accelerometers were Dytran Instruments, Inc., 3200BST shock accelerometers with a maximum 

rating of 10 000 g, and a frequency range of 0.35 to 10 000 Hz. The locations and directions of the 

accelerometers are documented in Section 3.4. 

Two sets of plots are provided, six for each shield to show the accelerations recorded during the test. 

Approximately 0.11 s of data was recorded at 100 000 samples per second for each channel. There is a 

significant difference in the level of the signals for the 20-ply panel and the 24-ply panel. In the test, the 

released blade impacted the frame of the 20-ply panel after penetrating the shield. This resulted in very 

high measurements on the 20-ply accelerometers and clipping on two of the channels. The 24-ply shield 

was not penetrated. The raw accelerometer data for the 12 accelerometers are shown in Figures 59 and 60 

for the 20- and 24-ply panels, respectively. The acceleration response for the 12 channels begins with a 

low value prior to detonation of the charge releasing the two blades. After detonation there is an increase 

in response, likely due to the pressure pulse from the detonation. Impact of the blade occurs at 

approximately 0.018 s, when the acceleration increases dramatically and then gradually decreases to a low 

level at approximately 0.1 s. 

The 20-ply frame data were compromised by the impact of the blade into the frame; therefore, no 

further processing was performed. For the reader’s convenience, power spectral density (PSD) plots are 

provided for the 24-ply-frame accelerometers in Figure 61. There was measurable frequency content in all 

the signals up to approximately 25 KHz, well below the Nyquist frequency. Above that the signal 

magnitudes were low. Data are shown for the frequency range of 0 to 1000 Hz. The PSDs were calculated 

with MATLAB (The MathWorks, Inc.) using the script in Figure 62. 

DIC was available during the test. In order to define the stand movement, the displacement was 

computed at a number of points on the perimeter of the 24-ply panel. The locations of points where 

displacements were measured are shown in Figure 63. The circumferential, axial, and radial 

displacements are provided for points 10 to 13 in Figures 64 to 67 for the four corners of the shield. The 

positive directions of axial and circumferential displacements are shown in Figure 63. Positive values of 

radial displacements are out of the page. In the plots, time 0 is the time of the first impact of the blade on 

the panel. This corresponds to a time of approximately 0.018 s in Figure 60. These displacements are 

located very close to the accelerometer locations. Computations were only possible up to 0.02 s after 

impact because of severe vibrations at the camera location making it impossible for the ARAMIS 

(GOM mbH, Germany) software to accurately track the pattern on the panel. Maximum displacement of 

the panel occurred at approximately 0.0065 s after impact, as shown in Figure 45. 
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References 

1. Carney, Kelly, et al.: Weight Assessment for Fuselage Shielding on Aircraft With Open-Rotor 

Engines and Composite Blade Loss. NASA/TM—2013-216582 (DOT/FAA/TC–13/34), 2013. 

2. Sweetman, Bill: The Short, Happy Life of the Prop-fan. Air & Space Magazine. 

2005. http://www.airspacemag.com/history-of-flight/the-short-happy-life-of-the-prop-fan7856180/#ixzz33gvo7PuE Accessed June 10, 2015. 

3. Bowman, C.L., et al.: Mechanical Properties of Triaxial Braided Carbon/Epoxy Composites. NASA 

Document ID 20040112004, 2003. http://ntrs.nasa.gov