Impact Testing of Composites for Aircraft

Engine Fan Cases

NASA/TM—2001-210887

April 2001

AIAA–2001–1633

Gary D. Roberts and Duane M. Revilock

Glenn Research Center, Cleveland, Ohio

Wieslaw K. Binienda and Walter Z. Nie

University of Akron, Akron, Ohio

S. Ben Mackenzie and Kevin B. Todd

Saint-Gobain Performance Plastics, Ravenna, Ohio
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Prepared for the

42nd Structures, Structural Dynamics, and Materials Conference and Exhibit

cosponsored by the AIAA, ASME, ASCE, AHS, and ACS

Seattle, Washington, April 16–19, 2001

Gary D. Roberts and Duane M. Revilock

Glenn Research Center, Cleveland, Ohio

Wieslaw K. Binienda and Walter Z. Nie

University of Akron, Akron, Ohio

S. Ben Mackenzie and Kevin B. Todd

Saint-Gobain Performance Plastics, Ravenna, Ohio

Impact Testing of Composites for Aircraft

Engine Fan Cases

NASA/TM—2001-210887 AIAA–2001–1633
Available from

NASA Center for Aerospace Information

7121 Standard Drive

Hanover, MD 21076

National Technical Information Service

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Available electronically at http://gltrs.grc.nasa.gov/GLTRS
NASA/TM—2001-210887 1 

IMPACT TESTING OF COMPOSITES FOR AIRCRAFT ENGINE FAN CASES 

Gary D. Roberts and Duane M. Revilock 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Wieslaw K. Binienda and Walter Z. Nie 

University of Akron 

Akron, Ohio 44325 

S. Ben Mackenzie and Kevin B. Todd 

Saint-Gobain (formerly Norton) Performance Plastics 

Ravenna, Ohio 

ABSTRACT

Before composite materials can be considered for use in 

the fan case of a commercial jet engine, the 

performance of a composite structure under blade-out 

loads needs to be demonstrated. The objective of this 

program is to develop an efficient test and analysis 

method for evaluating potential composite case 

concepts. Ballistic impact tests were performed on 

laminated glass/epoxy composites in order to identify 

potential failure modes and to provide data for analysis. 

Flat 7′′X7′′ panels were impacted with cylindrical 

titanium projectiles, and 15′′ diameter half-rings were 

impacted with wedge-shaped titanium projectiles. 

Composite failure involved local fiber fracture as well 

as tearing and delamination on a larger scale. A 36′′

diameter full-ring subcomponent was proposed for 

larger scale testing. Explicit, transient, finite element 

analyses were used to evaluate impact dynamics and 

subsequent global deformation for the proposed fullring subcomponent test. Analyses on half-ring and 

quarter ring configurations indicated that less expensive 

smaller scale tests could be used to screen potential 

composite concepts when evaluation of local impact 

damage is the primary concern. 

INTRODUCTION

High-bypass-ratio turbofan engines are used to power 

modern large commercial aircraft because of their high 

overall efficiency, high thrust at low flight speeds, low 

jet velocity, and low fuel consumption.1

 The fan case is 

the largest structural component in these engines. 

Currently, metal alloys are used for the case material in 

all commercial engines. Use of composite materials 

could significantly reduce the weight of the fan case, 

however technical feasibility and cost effectiveness first 

need to be demonstrated. The fan case in a jet engine is 

required to contain a fan blade in the rare event of a 

blade loss during engine operation. Because of the 

safety issues involved, full-scale demonstration of any 

new design is required by government regulations 

before the design can be used in commercial engines. 

This is an expensive process that has limited the 

development and use of new fan case materials and 

designs. Manufacturing cost has been a major limiting 

factor in the development and use of large composite 

structures for jet engines. However, fiber costs are 

projected to decrease substantially, and rapid 

developments are occurring in the area of low cost 

manufacturing of large composite structures. As a 

result, fabrication of a composite fan case should be 

economically feasible in the near future. Design of a 

composite case for normal engine operating conditions 

should also be technically feasible. However, the 

requirement for blade containment poses a unique 

challenge for the design of a composite fan case. 

There are two major technical challenges for the use of 

composite materials in a fan case. The first is to 

demonstrate that a containment system with a 

composite case can stop a failed blade from reaching 

the engine nacelle. The second is to demonstrate that 

the case can maintain structural integrity after a bladeout incident. Structural integrity is required to limit 

secondary damage caused by impact debris and to 

constrain out-of-balance motion of the rotor after a 

blade or blade fragment is lost. Ultimately, the 

performance of a composite fan case must be 

demonstrated in a full-scale engine blade-out test 

according to Federal Aviation Regulations.2

 In this test, 

a fan blade is released at full operating speed. Since the 

engine is destroyed in this test, it is usually the final test 

performed in the certification process for a new engine. 

Any changes to the design of the fan case at this stage 

can significantly increase cost and delay certification of 

the engine. These economic and schedule issues create 

tremendous pressure on the engine manufacturer to use 
NASA/TM—2001-210887 2 

traditional materials and designs. However, as engine 

design moves toward higher bypass ratios, there will be 

an increasing need to consider composite materials for 

the larger diameter fan cases. 

Explicit finite element analysis is beginning to be used 

earlier in the design cycle for metal fan cases to shorten 

development time, reduce testing cost, and reduce the 

risk of failure for the final design. This approach is also 

used in the automobile industry where explicit finite 

element analysis is used in conjunction with 

instrumented crash tests to improve vehicle 

crashworthiness. Recent advances in the use of explicit 

codes such as LsDyna3D and ABAQUS/Explicit allow 

numerical testing of an entire car before even a single 

car body is manufactured. A similar approach is needed 

to overcome the large development cost and technical 

risk associated with using a new composite fan case 

design. Use of finite element analysis to assist design is 

most effective when a design methodology already 

exists and a large database (laboratory tests, 

certification tests, and field service) is available for 

validation. This is a particularly important issue for 

composites because of the large number of possibilities 

for fiber orientation, failure mode, materials, and 

manufacturing method. Unfortunately, there is no prior 

certification or service experience for a composite case, 

and only a limited amount of laboratory data is 

available from government and commercial technology 

development programs. 

The two main goals of this work are: (1) to develop and 

demonstrate design concepts for a composite fan case 

and (2) to generate test data and analysis methods for 

future use as part of an overall design methodology. 

The first step is to evaluate the capabilities and 

limitations of current materials, test methods, and 

explicit finite element analysis methods. In this paper, 

ballistic impact test results are reported for composite 

panels. A larger full-ring structure is proposed as a 

possible subcomponent design. Explicit finite element 

analysis is used to evaluate impact dynamics for the 

proposed subcomponent test. 

COMPOSITE IMPACT TESTS

Ballistic impact tests have been used on targets ranging 

in size from small flat panels to full-scale 

subcomponents. Full-scale tests can simulate many of 

the conditions in a real blade-out incident. Small panel 

tests attempt to simulate at least a few critical 

parameters. The impact velocity is usually chosen as 

the center of mass velocity of the blade fragment. The 

projectile shape and orientation relative to the target is 

determined from the blade dynamics and the type of 

damage observed in the case. This information must be 

obtained from full-scale testing and analysis since it 

depends on the particular engine design and operating 

conditions. The design of the fan containment system in 

a particular engine depends on issues such as weight, 

cost, prior experience, and limitations on space 

available for case deflection. Two design approaches 

are currently used for blade containment. In the 

“hardwall” design, the wall thickness near the blade 

impact site is made sufficiently thick to resist 

penetration, and ribs are used to control global stiffness. 

In this design, local damage at the impact site must be 

minimized, and propagation of the damage must be 

limited so that global stiffness and strength of the 

structure are maintained. In the “softwall” design, the 

metal case is penetrated by the blade and captured by an 

outer wrap of fabric. In this design, stiffness is provided 

by either an isogrid or honeycomb structure on the case. 

Although the blade is allowed to penetrate the case in 

the softwall design, it is critical that the perforated area 

be kept as small as possible and that the damaged area 

does not grow to affect global stiffness and strength. 

The high specific stiffness and low density of 

composites are advantages for both hardwall and 

softwall designs, but poor impact resistance is a 

potential problem. Low velocity impact of composites 

can cause extensive internal damage (microcracking, 

delamination) with little visible surface damage.3

 High 

velocity tests on composites using flat, hemispherical, 

and conical projectiles have shown shear failure in plies 

nearest the impact surface to be the initial failure 

mode.4

 Subsequent damage modes depend on the 

materials, fiber architecture, and ply lay-ups, but often 

include delamination and fiber tensile failure in the 

interior or near the back surface plies. 

In order to identify expected damage modes under 

simulated blade-out conditions, a 15′′ diameter, 0.420′′

thick 2D laminated glass/epoxy half ring was impacted 

with a wedge-shaped titanium projectile at a velocity of 

850 ft/s. The 0.73lb (330g) wedge was 5′′ long by 3.75′′

wide with the thickness increasing from 1/8′′ at the thin 

edge to 3/8′′ at the thick edge to simulate the mass 

distribution of a fan blade fragment. To simulate the 

blade dynamics, the desired orientation of the wedge 

was 45° from the horizontal with the thin edge of the 

wedge making contact first. The thin edge curls after 

contact as the wedge rotates and the thick edge makes a 

second contact. Figure 1 shows the wedge and a side 

view of the 15′′ half-ring after impact. Local fiber shear 

fracture occurs along the lines where the thin and the 

thick edges of the wedge make contact. Extensive 

delamination occurs throughout the composite, and 

tearing occurs in plies near the back surface. Damage 

also appears to be concentrated at some points around 
NASA/TM—2001-210887 3 

the edges where the panel was held in place by the 

metal fixtures. 

The damage evident in Figure 1 indicates that a 

composite hardwall containment system will require 

use of a surface layer that is capable of eliminating 

local shear failure. In addition, some method for 

controlling delamination will be needed. Lightweight 

materials and structures that are capable of absorbing a 

large amount of energy have been developed for armor 

applications. Although most of these materials are not 

suitable for use as the main structural material in a 

containment case, a recent report5

 has examined the 

feasibility of applying some of these armor concepts to 

commercial aircraft in the form of protective barriers 

that could enhance aircraft survivability in the event of 

an uncontained engine failure. One such concept is to 

place a ceramic plate on the impact surface of the 

composite. The ceramic serves to fragment the 

projectile and spread the impact load over a larger area 

of the backing plate. Projectile fragmentation occurs at 

very high impact velocities as a result of shock waves 

within the projectile. This process is unlikely to occur 

with large titanium blade fragments impacting a case at 

velocities characteristic of a blade-out event. However, 

a ceramic layer may still be useful for spreading the 

impact load and blunting sharp blade fragments. 

Delamination can be controlled by reinforcement in the 

transverse direction by processes such as stitching, 

pinning, or 3D weaving. Evaluation of these concepts 

using the test configuration shown in Figure 1 would be 

expensive and time consuming. The high cost is a result 

of fabrication costs for the panels, fabrication costs for 

the projectile, and the cost of performing the test. 

Although the proposed concepts would not necessarily 

be too expensive for commercial production, 

fabrication of a small number of test panels can be very 

costly. For example, set-up costs for weaving represent 

a small part of the total cost for a large production run 

but a large part of the cost for fabrication of a small 

number of panels. In addition, scheduling such a small 

run often cannot be done in a timely manner. The 

wedge used for the test shown in Figure 1 requires 

machining of a block of titanium. A simpler projectile 

that could be cut from stock material would be less 

expensive and more readily available. In addition, the 

impact test shown in Figure 1 is somewhat complex. 

Since the projectile is not cylindrical or spherical, it 

must be mounted in a cylindrical container (sabot) in 

order to fit into the barrel of the gas gun. The 45° pitch 

angle is achieved by bonding the titanium wedge to a 

45° prism-shaped wood platform within the sabot. 

Imprints of the wedge in the test panel and high-speed 

video of the projectile during the test have shown that 

the orientation of the wedge often deviates in pitch and 

roll from the desired orientation. Deviations in the 

location of impact can also cause variability in the 

results. The location and orientation of the blade has a 

large effect on the damage induce by impact. Damage 

induced by the first contact of the thin edge of the 

wedge is more severe for pitch angles less than 45°. 

Rotation of the wedge also causes more severe damage 

because the wedge makes contact at its corners rather 

than along its edges. All of these factors indicate that a 

simpler, less expensive test would be preferable for 

preliminary screening of composite concepts. In this 

paper, a simpler test was used to evaluate the 

performance of a ceramic layer on a glass/epoxy 

composite and the performance of 3D woven materials. 

Some justification based on observed failure modes is 

discussed below. Analysis of impact dynamics for a 

full-ring structure, discussed in a later section, provides 

further support for the use of the simpler test method. 

Small (7′′X7′′) flat panels were impacted with flatended cylindrical titanium projectiles (2.54 cm long, 

1.27 cm diameter). Panel and projectile shapes are 

shown in Figure 2. This test is less realistic than tests 

with the wedge-shaped projectile in the sense that it 

cannot simulate blade dynamics. However, preliminary 

tests on laminated composite panels with various fiber 

and matrix materials demonstrated that the impact 

damage had features similar to those shown in Figure 1. 

Fiber shear fracture occurred on the front surface, 

delamination occurred in the interior, and fiber tearing 

occurred on the back surface. The ballistic limit was 

determined by testing approximately 6-8 panels over a 

range of impact velocities. Ideally, the ballistic limit is 

the velocity below which the projectile does not 

perforate the panel and above which the projectile 

completely perforates the panel. The ballistic limit can 

then be defined as the average of the highest velocity 

that did not perforate and the lowest velocity that 

completely perforated the panels. In practice, there may 

be a small velocity region near the ballistic limit in 

which non-perforation and perforation overlap. In 

addition, failure may be defined by a criterion other 

than complete perforation. For example, a plate may be 

considered to fail the test if a crack propagates through 

the panel, even though the projectile does not perforate 

the panel. The above definition for ballistic limit can 

still be used, as long as the failure criterion is defined. 

Impact test data for five different types of composite 

panels are shown in Figure 3. For these composite 

panels there was a velocity region near the ballistic 

limit within which the projectile was caught in the 

panel. This was interpreted as perforation (or failure) 

for the purpose of calculating the ballistic limit. The 

ballistic limits calculated from the data in Figure 3 are 

shown in Figure 4 along with the ballistic limits that 

have been measured for 304 stainless steel and for 
NASA/TM—2001-210887 4 

Inconel 718. The stainless steel alloy was chosen for 

comparison because it has been used in commercial fan 

cases. The ballistic limits of Inconel 718 in the annealed 

and heat treated forms are shown to indicate the range 

of properties possible for some alloys by heat treatment. 

The thickness (or areal weight) range of the metal alloy 

panels used to obtain the data in Figure 4 was chosen so 

that the ballistic limits would fall within the range of 

velocities which are characteristic of blade fragments 

during a blade-out. It is useful to have data over a range 

of areal weights because it is usually not feasible to 

fabricate panels of different materials with exactly 

equal areal weights for direct comparison. 

The first composite concept evaluated was the use of a 

1.9mm (0.075′′) thick silicon carbide ceramic layer to 

reduce the extent of shear failure at the impact surface. 

7′′X7′′ glass/epoxy panels were cut from a larger 

16.5′′X38.0′′ panel that was made by hand lay-up of 28 

plies of type 6781 S2 glass/E-761 epoxy prepreg with 

autoclave cure. Six of the 7′′X7′′ panels were used to 

measure the ballistic limit of the composite (Figure 3). 

In Figure 4, the ballistic limit of the S2 glass/E-761 

composite is 191m/s (626ft/s), which is well below that 

of the metal alloys at an equivalent areal weight. 

Figures 2a and 2b show the front and back surfaces 

after impact for glass/epoxy panels with and without a 

ceramic front layer. The lower panel in Figure 2a has a 

nylon spall sheet bonded to the surface to catch ceramic 

fragments during the test. The ceramic fragments have 

been removed to view the surface of the composite. The 

panels in Figure 2 were impacted at 636ft/s (194m/s), 

just above the ballistic limit. The panel without the 

ceramic layer exhibited the typical failure modes of 

fiber shear failure on the front, delamination in the 

middle plies, and tearing on the back. The ceramic layer 

completely eliminated the fiber shear fracture on the 

front surface but did not reduce the extent of 

delamination as viewed from the back side of the 

panels. The dashed circles in Figure 2b show the 

delamination areas visible from the back side of the 

panels. (Delamination is clearly visible when viewing 

the panels but does not show well in the photographs). 

Although the ceramic layer is effective in eliminating 

fiber shear failure at the impact surface, this benefit 

comes at the expense of increasing the areal weight of 

the panel. Further tests are needed to determine the 

optimum ceramic thickness and the effectiveness of the 

ceramic on a total areal weight basis. 

The second concept evaluated was use of a 3D woven 

architecture to reduce delamination. Composites were 

fabricated using 3D carbon fiber preforms and resin 

transfer molding. The main system of interest was IM7 

fiber with the toughened epoxy resin, PR520. However, 

options were also considered for potential cost 

reduction. For material cost reduction, E-862 epoxy 

resin was used in place of PR520, and AS4 fiber was 

used in place of IM7. For weaving cost reduction, tows 

of the AS4 fiber were combined during weaving. This 

resulted in a courser weave for the AS4/PR520 and 

AS4/E-862 composites compared to the IM7/PR520 

and IM7/E-862 composites. Failure on the front surface 

of the 3D woven composites was similar to that of the 

2D glass/epoxy laminated composites with local shear 

failure on the front surface. However, the subsequent 

damage pattern was different. For the 3D composites 

there was no evidence of delamination, but there was a 

region of microcracking around the impact site, 

particularly for the more brittle E-862 resin. The 3D 

composites had fiber breakage but no ply tearing on the 

back side. The IM7/PR520 composite had a ballistic 

limit of 190m/s (623ft/s), comparable to that of the 

heat-treated Inconel 718 at an equivalent areal weight. 

The ballistic limit of the IM7/E-862 composite was 

lower at 179m/s (587ft/s). Ballistic limits for 

composites made using the courser weave and AS4 

fibers were 156m/s (512ft/s) for AS4/PR520 and 

159m/s (522ft/s) for AS4/E-862. This is much lower 

than the ballistic limits of the metal alloys in Figure 4. 

The IM7 composites exhibited a slight permanent 

conical deformation in the impact region when tested 

near the ballistic limit. This type of permanent 

deformation is typical for metal alloys but not for 

composites. This effect occurred to a much smaller 

extent in the AS4 composites and not at all in the 2D 

laminated glass/epoxy composites. 

SUBCOMPONENT DEFINITION

Although more small panel testing is needed to 

optimize performance, there are three results from the 

small panel tests that indicate scale-up is justified. First, 

the ballistic limit of some composites approaches that 

of the metals, even though the materials and fiber 

architecture have not yet been optimized. Second, 

transverse fiber reinforcement is effective in limiting 

delamination. Third, the ceramic layer is effective in 

eliminating fiber shear failure at the impact surface. 

The next level of testing should address the issues of 

manufacturability and cost as well as demonstrating 

technical feasibility at a larger scale. The subcomponent 

test in Figure 5 is proposed for this purpose. The 36′′

diameter ring will be manufactured by filament 

winding. Filament winding is a low cost process for 

fabricating cylindrical structures. This process also 

allows for considerable flexibility in materials selection 

and fiber architecture. In addition, features such as ribs 

and flanges can be included in the design. If the 

composite material properties are not strain rate 

dependent, scaling of displacement, contact forces, and 

strains from a small test to a larger prototype is 
NASA/TM—2001-210887 5 

possible, but scaling of damage is not.6

 Therefore, 

results of the small panel tests can be used as a guide to 

the design of the subcomponent in Figure 5, but some 

preliminary testing and development of a numerical 

model will be needed to implement a cost effective test 

program at this larger scale. 

FINITE ELEMENT ANALYSIS

Explicit, transient, finite element analysis of the 

proposed subcomponent in Figure 5 was performed 

using the commercial code ABAQUS/Explicit.7

 A 

preliminary analysis was performed to examine the 

overall dynamics of the ring during and after impact. In 

this analysis the impact velocity was 600ft/s and the 

composite ring was taken to be 0.3′′ thick with the 

following properties, which are representative of a 

glass/epoxy laminate with a 0°/90° layup. 

E11= E22= 2.5X106

 lbf/in2

G12= 1.0X106

 lbf/in2

G13= G23= 6.0X105

 lbf/in2

υ12= υ21= 0.14 

Linear elastic, rate independent properties were used, 

and damage was not considered for this preliminary 

analysis. The projectile was a flat circular cylinder with 

diameter of 4.5′′, thickness of 0.75′′, and mass of 1.9lb. 

It was modeled as a rigid body with a density of 

0.0004086lbf2

/in4. The ring was inclined at an angle of 

22° to the horizontal. The mesh in Figure 6 along with 

the symmetry condition about the mid plane was used 

for the full-ring analysis. Half-ring and quarter-ring 

structures with both free and fixed boundary conditions 

were also analyzed. Four-node thin shell S4R elements 

were used for the shell. Six-node and eight-node solid 

C3D6 and C3D8R elements were used for the 

projectile. 

Deformation of the full ring structure at various time 

intervals is shown in Figure 7. At 0.11ms the case 

begins to bulge as the top edge of the projectile pushes 

on the case. At 0.31ms the center of the case bulges as 

the center of mass of the projectile contacts and pushes 

on the case. At later times the projectile rebounds, and a 

flexural wave begins to travel around the case. The 

displacement and velocity of the upper point on the 

projectile are shown in Figures 8a and 8b for the fullring model along with the results for the half- and 

quarter-models with both free and fixed boundary 

conditions. In the time interval from 0 to 0.25ms, 

contact between the upper point on the projectile and 

the case causes the case to bulge and causes the 

projectile to slow down and rotate. Between .25ms and 

0.7ms the center point of the projectile contacts the case 

and causes the case to bulge further. This causes the 

velocity of the upper point on the projectile to increase 

during this time interval. Between 0.7ms and 1.0ms the 

projectile rebounds from the case. After 1.0ms the 

projectile releases from the case and continues to travel 

at constant velocity. In the time interval from 0 to 

1.0ms, while the projectile is in contact with the case, 

the motion of the projectile is very similar for the fullring and the half-rings with both free and fixed 

boundary conditions. The motion of the projectile 

impacting the quarter-rings deviates from the half- and 

full-ring configurations after 0.4ms when the initial 

impact of the upper point on the projectile with the case 

is complete. The same conclusions result when 

displacements and velocities for points on the rings are 

analyzed. These results suggest that a half-ring 

configuration could be used in place of a full ring to 

evaluate local damage. Either free or fixed boundary 

conditions could be used. Alternatively, two impact 

tests could be performed 180° apart on the same fullring. Either approach would reduce test costs. A 

quarter-ring configuration could be used if the purpose 

of the test is to evaluate local damage caused by a 

single impact with no subsequent complex motion of 

the projectile. This supports the use of a simple 

projectile and smaller panels for initial screening 

studies when local damage is the primary concern. 

Strain energies stored in arcs of various sizes around 

the impact point are shown in Figure 9a for the full-ring 

configuration. The total energy (strain energy + kinetic 

energy) is shown in Figure 9b. An additional curve 

showing the energy in the full-ring plus the energy in 

the projectile is shown in each figure. Since the 

projectile is modeled as a rigid body, all of its energy is 

in the form of kinetic energy. In Figure 9a the strain 

energy grows to a maximum value at about 0.45ms as 

the projectile comes into contact with the ring. The 

maximum strain energy is a little over half of the initial 

kinetic energy of the projectile. In Figure 9b the total 

energy in the full-ring remains constant after about 

1.0ms because the projectile has rebounded from the 

ring. The maximum strain energy density in the ring is 

a useful parameter for preliminary design of the ring. 

From Figure 9a it appears that the maximum strain 

energy density would occur at about 0.40ms within a 

50° arc. This energy density could be used in a static 

composite analysis to determine how close the ring is to 

its ultimate strength. A static analysis could also be 

used to suggest changes in materials, fiber architecture, 

or other parameters that would improve performance. 

The static analysis could also be used interactively with 

the filament winding design and machine control 

software to identify structures that will perform well but 

also be feasible to fabricate. Future efforts will focus on 

refining the finite element methods to model damage 

more precisely and include strain rate dependence. 
NASA/TM—2001-210887 6 

Efforts will also continue to refine the subcomponent 

design. 

CONCLUSIONS

Preliminary explicit finite element analyses suggest that 

small panel tests can be a cost effective way to evaluate 

local impact resistance of a composite at the first point 

of contact with the projectile. Results of impact tests on 

7′′X7′′ panels indicate that the ballistic limit of 

composites can approach that of metals. A protective 

ceramic surface layer is a promising approach for 

preventing fiber shear failure on the impact surface but 

not for limiting delamination. Transverse fiber 

reinforcement can be used to control delamination 

during impact. Use of a proposed 36′′ diameter full-ring 

subcomponent manufactured by filament winding is a 

reasonable approach for assessing both technical and 

economic feasibility of composite concepts at a larger 

scale than the small panel tests. Small panel testing, 

finite element analysis, static analysis, and preliminary 

design analysis using filament winding software can be 

integrated to reduce the time and expense of testing 

larger subcomponents. 

REFERENCES

1. Jack D. Mattingly, “Elements of Gas Turbine 

Propulsion”, pp xlviii-xlix, McGraw-Hill, Inc., 

1996. 

2. Federal Aviation Regulations Part 33.94. U.S. 

Federal Aviation Administration 

3. Greszczuk, L.B., “Damage in Composite Materials 

due to Low Velocity Impact”, Impact Dynamics, 

Chapter 3, John Wiley & Sons, Inc. (1982), p93. 

4. R.A.W. Mines, A.M. Roach, and N. Jones, “High 

Velocity Perforation of Polymer Composite 

Laminates”, Int. J. Impact Eng. 22 (1999) 561-588. 

5. U.S. Department of Transportation/Federal 

Aviation Administration Report No. 

DOT/FAA/AR-97/53, September 1997. 

6. Abrate, Serge, “Impact on Composite Structures”, 

pp132-4, Cambridge University Press, 1998. 7. ABAQUS/ Explicit 5.8, Habbitt, Karlsson & 

Sorensen, Inc. 
NASA/TM—2001-210887 7 

Figure 1. Glass/epoxy 15” diameter half-ring and wedge-shaped titanium projectile 

(a) Front surface (b) Back surface 

Figure 2. Glass/epoxy 7”X7” panels and cylindrical titanium projectile. Bottom panel had a ceramic layer under 

 a nylon spall sheet on the front surface. Ceramic is removed to view the surface of the composite. 

Delamination 

Delamination 
NASA/TM—2001-210887 8 

Figure 3. Composite impact test data from 7”X7” flat panel tests. 

Figure 4. Ballistic limit for various composites and metal alloys.

Fiber/Resin

IM7/PR520 IM7/E-862 AS4/PR520 AS4/E-862 Glass/Epoxy

Impact Velocity, m/s

150

160

170

180

190

200

210

Not Perforated

Caught

Perforated

Areal Weight, kg/m2

6 8 10 12 14 16 18 20 22 24

Ballistic Limit, m/s

140

160

180

200

220

240

260

280

300

S2 Glass/E-761

Inconel 718, Annealed

Inconel 718, Heat Treated

IM7/PR520

IM7/E-862

AS4/PR520

AS4/E-862

304 Stainless Steel
NASA/TM—2001-210887 9 

Figure 5. 36” diameter full-ring subcomponent test. 

Figure 6. Finite element mesh for 36” diameter full-ring subcomponent. 

Figure 7. Deformation of 36” full-ring subcomponent at various times after impact.

1" thick steel fixture 

Composite ring 

22°

36” 

34” 

6”

12" 

Titanium disk 

(4.5" dia., 0.75" thick) 

t= 0.02ms t= 0.31ms t= 0.11ms 

t= 0.80ms t= 1.00ms t= 2.50ms 
NASA/TM—2001-210887 10 

Time, ms

Velocity, in/s

(a) Displacement of upper point on the projectile

(b) Velocity of upper point on the projectile

Figure 8. Finite element analysis results for a titanium disk projectile impacting full-ring, half-ring and

quarter-ring composite structures at 600ft/s. Free and fixed boundary conditions (BC) were used for

half- and quarter-rings. (a) Displacement and (b) velocity of the upper point on the projectile.

0

2,000

-4,000

4,000

-2,000

Time, ms

0.0 1.0 2.0 3.0 

Displacement, in

-10.0

-5.0

0.0

Full-ring

Half-ring

Fixed BC

Half-ring

Free BC

Quarter-ring

Fixed BC

Quarter-ring

Free BC

0.0 1.0 2.0 3.0 

Full-ring

Half-ring

Fixed BC

Half-ring

Free BC

Quarter-ring

Fixed BC

Quarter-ring

Free BC
NASA/TM—2001-210887 11 Strain Energy, lbf-in Total Energy, lbf-in

(b) Total energy in various arc sections centered on the point of impact

Figure 9. Finite element analysis results for a titanium disk projectile impacting full-ring composite structures at

600ft/s. Energy stored in arc sections centered on the point of impact is indicated for various arc

angles. (a) Strain energy and (b) Total energy (strain energy + kinetic energy)

(a) Strain energy in various arc sections centered on the point of impact

20°

50°

Full-ring

Full-ring and

Projectile

(same as full-ring)

180°

90°

20°

50°

Full-ring

Full-ring and

Projectile

180°

90°

Time, ms

0.0 1.0 2.0 3.0 

0

30,000

10,000

35,000

15,000

20,000

25,000

5,000

Time, ms

0.0 1.0 2.0 3.0 

0

50,000

10,000

60,000

20,000

30,000

40,000
This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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AIAA–2001–1633

E–12763

WU–708–24–13–00

17

Composites; Polymer matrix composites; Impact tests

Unclassified -Unlimited

Subject Category: 24 Distribution: Nonstandard

Prepared for the 42nd Structures, Structural Dynamics, and Materials Conference and Exhibit cosponsored by the AIAA, ASME, ASCE,

AHS, and ACS, Seattle, Washington, April 16–19, 2001. Gary D. Roberts and Duane M. Revilock, NASA Glenn Research Center;

Wieslaw K. Binienda and Walter Z. Nie, University of Akron, 302 Buchtel Mall, Akron, Ohio 44325–0001; S. Ben Mackenzie and

Kevin B. Todd, Saint-Gobain Performance Plastics, Ravenna, Ohio (formerly Norton). Responsible person, Gary D. Roberts,

organization code 5150, 216–433–3244.

Gary D. Roberts, Duane M. Revilock, Wieslaw K. Binienda,

Walter Z. Nie, S. Ben Mackenzie, and Kevin B. Todd

Impact Testing of Composites for Aircraft Engine Fan Cases

Before composite materials can be considered for use in the fan case of a commercial jet engine, the performance of a

composite structure under blade-out loads needs to be demonstrated. The objective of this program is to develop an

efficient test and analysis method for evaluating potential composite case concepts. Ballistic impact tests were performed

on laminated glass/epoxy composites in order to identify potential failure modes and to provide data for analysis. Flat

7×7 in. panels were impacted with cylindrical titanium projectiles, and 15 in. diameter half-rings were impacted with

wedge-shaped titanium projectiles. Composite failure involved local fiber fracture as well as tearing and delamination on

a larger scale. A 36 in. diameter full-ring subcomponent was proposed for larger scale testing. Explicit, transient, finite

element analyses were used to evaluate impact dynamics and subsequent global deformation for the proposed full-ring

subcomponent test. Analyses on half-ring and quarter ring configurations indicated that less expensive smaller scale tests

could be used to screen potential composite concepts when evaluation of local impact damage is the primary concern.