Sandi G. Miller, Katherine M. Handschuh, Matthew J. Sinnott, Lee W. Kohlman, 

Gary D. Roberts, Richard E. Martin, Charles R. Ruggeri, and J. Michael Pereira

Glenn Research Center, Cleveland, Ohio

Materials, Manufacturing, and Test Development of a 

Composite Fan Blade Leading Edge Subcomponent

for Improved Impact Resistance

NASA/TM—2015-218340

February 2015
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Sandi G. Miller, Katherine M. Handschuh, Matthew J. Sinnott, Lee W. Kohlman, 

Gary D. Roberts, Richard E. Martin, Charles R. Ruggeri, and J. Michael Pereira

Glenn Research Center, Cleveland, Ohio

Materials, Manufacturing, and Test Development of a 

Composite Fan Blade Leading Edge Subcomponent

for Improved Impact Resistance

NASA/TM—2015-218340

February 2015

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Materials, Manufacturing, and Test Development of a 

Composite Fan Blade Leading Edge Subcomponent 

for Improved Impact Resistance 

Sandi G. Miller, Katherine M. Handschuh,* Matthew J. Sinnott,* Lee W. Kohlman, Gary D. Roberts, 

Richard E. Martin, Charles R. Ruggeri, and J. Michael Pereira 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Abstract 

Application of polymer matrix composite materials for jet engine fan blades is becoming attractive as 

an alternative to metallic blades; particularly for large engines where significant weight savings are 

recognized on moving to a composite structure. However, the weight benefit of the composite is offset by 

a reduction of aerodynamic efficiency resulting from a necessary increase in blade thickness; relative to 

the titanium blades. Blade dimensions are largely driven by resistance to damage on bird strike. Further 

development of the composite material is necessary to allow composite blade designs to approximate the 

dimensions of a metallic fan blade. The reduction in thickness over the state of the art composite blades is 

expected to translate into structural weight reduction, improved aerodynamic efficiency, and therefore 

reduced fuel consumption. This paper presents test article design, subcomponent blade leading edge 

fabrication, test method development, and initial results from ballistic impact of a gelatin projectile on the 

leading edge of composite fan blades. The simplified test article geometry was developed to realistically 

simulate a blade leading edge while decreasing fabrication complexity. Impact data is presented on 

baseline composite blades and toughened blades; where a considerable improvement to impact resistance 

was recorded. 

1.0 Introduction 

The purpose of the fan blades on high bypass turbofan jet engines is to accelerate air for propulsion, 

while a smaller fraction is passed into the compressor section of the engine. At take-off and landing, these 

fan blades are susceptible to potentially damaging bird strike events that could lead to engine shutdown. 

As a result, blade impact strength specifications are outlined by the Federal Aviation Administration 

(FAA) and are prerequisite to blade implementation into commercial aircraft (Ref. 1). Meeting the 

standards set forth by the FAA requires a necessary, but costly, full scale engine test where the pass/fail 

criteria are dependent on engine size. FAA regulations outline three categories of bird strike: large single 

bird, small or medium flocking bird, and large flocking bird. For a large single bird, the ingestion speed 

of the FAA test of turbo fan engines currently in service is 200 kn (337.5 ft/s). It is understood that the 

engine will shut down with the ingestion of a large single bird. Following ingestion of small and medium 

flocking bird(s) the engine must generate 75 percent of take-off power or thrust and with only a 

momentary drop of less than 3 sec. For large flocking birds, the engine needs to operate for a 20 min ‘runon’. The test to meet these criteria includes evaluation of engine performance at different power ratings 

until shutdown. These standards affect the design parameters of fan blades. 

The primary drivers for technology development within the commercial aircraft community include 

cost savings and efficiency gains. Common methods to meet these challenges include weight reduction 

and increased aerodynamic efficiency of aircraft components. The large intake fan blades are an attractive 

technology for development as both weight reduction and enhanced efficiency are feasible. The

*

 NASA Lewis’ Educational and Research Collaborative Internship Project (LERCIP) Intern 

NASA/TM—2015-218340 1 

 
successful application of carbon fiber reinforced polymer matrix composite (PMC) fan blades by GE 

(Ref. 2) has significantly reduced engine weight; relative to a structure using titanium blades. The weight 

reduction is recognized not only in the blades themselves, but also in secondary engine components such 

as bearings, mounts, and fan containment structures. The limitation of the composite blades however, is 

that the PMC turbofan blades must be thicker than conventional titanium blades to meet the strength 

requirements for bird strikes set forth by the FAA. Thus, the aerodynamic efficiency of the PMC blades is 

reduced when compared to titanium blades. This paper addresses two parallel approaches to improving 

the aerodynamic efficiency of PMC blades, including (1) decreased fan blade thickness through improved 

damage tolerance of the composite, and (2) material studies that would enable camber change of the blade 

between cruise and take-off. The method considered to improve damage tolerance includes incorporation 

of a thermoplastic interlayer to enhance interlaminar strain capability, and reduce delamination on impact. 

The second path to increased aerodynamic efficiency of fan blades includes embedding shape memory 

alloy (SMA) ribbons into composite blade. Embedding such materials holds the possibility for variable 

camber fan blades; where shape change occurs between take-off and cruise speeds for increased fan 

efficiency. Currently fan blades are typically designed for cruise and operate at reduced efficiency during 

take-off. SMAs are metallic materials that undergo thermally and/or mechanically induced phase 

transitions (martensitic/austenitic) resulting in high reversible strain capability. These phase transitions, 

when thermally induced, can be used to generate stress and/or strain for actuation. 

The specific goals of this study were to (1) provide an impact test method for a blade leading edge 

subcomponent structure, (2) design a realistic laminate lay-up for the leading edge subcomponent test 

articles, and (3) identify material solutions to improve composite impact performance and allow for 

thinner fan blade structures. The test article design and test methodology were developed based on 

published experimental work bird strike simulation. 

2.0 Background 

2.1 Test Method Development 

References 3 and 4 summarize efforts where a simplified blade test article was used to evaluate the 

influence of variation in composite materials and ply configuration on impact resistance. The fan blades 

were simulated with a double tapered, diamond shape cross section with thickness decreasing from 

3.8 mm (0.15 in.) to 0.25 mm (0.01 in.) across a width of 7.6 cm (3 in.) and a length of 20.3 cm (8 in.). 

The simplified test blades were held as a horizontal cantilever for impact testing along the 7.6 cm (3 in.) 

width. Work in Reference 3 varied impact projectiles to include: ice, bird simulant, and steel spheres. The 

projectiles were aimed at the leading edge; which consisted of four pre-impregnated sheets of carbon fiber 

and epoxy. Conclusions from the testing in References 3 and 4 are presented below as background. 

2.2 Laminate Lay-Up Configuration 

Test data presented in References 3 and 4 provide a comparison between laminate lay-ups and these 

articles provided the basis for the leading edge design features implemented in this paper. 

The effects of ply configuration on material response to impact were compared in Reference 3. The 

two configurations compared included what was referred to as a standard [(±45°)2 | 0°10]s lay-up and a 

chordwise stiffened [(90°,0°)4 | 0°6]s lay-up with leading edge plies of –45°, +45°, +45°, –45° and 0°, 90°, 

90°, 0°, respectively. Based on ice and gelatin impact, the chordwise stiffened lay-up resisted impact 

damage to a greater extent than the standard lay-up, due to the higher stiffness of the 0° plies. 

Reference 4 describes three separate ply lay-ups; [(+45°/0°/–45°/0°)2/0°2]s, [(+22°/0°/–22°/0°)2/0°2]s, 

and [(+45°/0°/–45°/90°)2/0°2]s. In the blade impact tests, gelatin sphere projectiles of 2.54 cm (1 in.) 

diameter were shot at an angle of 30° to the blade chord at three different speeds: 90, 150, and 275 m/s. 

All test articles had the least damage at the lowest speed and each lay-up type encountered significant 

NASA/TM—2015-218340 2 
damage as speed increased. However, it was the third lay-up, which contained 90° plies that was deemed 

inferior as the loss of material was greater than either lay-up containing both ±45° and 0° ply orientations. 

2.3 Material 

References 3 and 4 were published in 1973 and 1974, respectively. Since that time, there have been 

advancements in composite materials that could potentially improve the performance of composite fan 

blades. Such material advancements include increased stiffness and strength of carbon fibers, toughened 

epoxy, and alternative laminate toughening approaches. The potential benefits of these material 

advancements for future blade design are considered in this paper. 

References 3, 4, and 5, compare material variations in flat panels and blade structures tested for 

impact resistance. Hybrid fiber materials along with strictly carbon fiber-based composites were 

evaluated through impact testing. Reference 3 reports a combination of Modmor II carbon fibers and 

PR-286 epoxy in combination with Kevlar and S-glass for improved transverse damage resistance. 

Reference 4 also evaluates Modmor II carbon fibers and various percentages of S-glass fibers for hybrid 

impact analysis with PR-286 epoxy. In these references the hybrid materials outperformed the carbon 

fiber baseline specimens in impact resistance, due to the higher strain to failure of the S-glass and Kevlar. 

For the particular application of thinner fan blades, the addition of fibers of higher strain to failure would 

benefit impact resistance, but with a stiffness penalty. Therefore, with the hybrid fiber approach, blade 

thickness would likely increase to meet the functional stiffness essential for loading. With a blade 

application in mind, this paper considers only carbon fiber based composites. 

In addition to constituent material consideration, each reference recommended application of leading 

edge protection to uniformly distribute impact load and decrease break-out at the leading edge on impact. 

To protect the double tapered leading edge specimen, References 3 and 4 utilized a stainless steel insert, 

bonded to the leading edge. A 0.15 cm (0.006 in.) thick stainless steel foil was then welded to the insert 

covering 2.5 cm (1 in.) of the specimen’s width. The leading edge protection helped eliminate break out, 

but delamination occurred between the composite and metal with larger projectile impacts. 

The literature review provided guidance for creating a baseline ply configuration and test method for 

a simplified leading edge test specimen. Insight to some impact resistant material and design 

combinations, as well as guidance toward leading edge protection, provided an initial test method and 

preliminary design reported in this paper. 

The overall goal of this effort was to define a test article, test method, and materials approach for 

evaluating thickness reduction of composite blades while meeting stiffness and impact requirements for 

safe operation. The reduction in blade thickness will enable optimization of aerodynamic efficiency over 

the current state of the art composite blade. The approach used in this paper was to provide a material 

interlayer capable of toughening the composite structure to yield acceptable impact resistance. A 

simplified test article geometry was required for repeatability and ease of manufacturing. A standardized 

test method was necessary for comparison of baseline and toughened test articles. Impact testing was 

performed to simulate realistic leading edge impact results and understand the effects. 

3.0 Experimental 

3.1 Materials 

IM7/8551-7 carbon fiber/epoxy pre-impregnated unidirectional sheets were purchased from Hexcel 

for leading edge fabrication. The manufactures recommended cure cycle is shown in Figure 1. The cured 

ply thickness is reported as 0.142 mm (0.0056 in.). The elastic moduli of the material are 158 GPa in the 

fiber direction and 8.3 GPa in the transverse fiber direction (Ref. 6). Based on experimental work, the 

Poisson’s ratios were determined from 0° and 90° tensile tests to be 0.333 and 0.012 respectively (Ref. 7). 

Due to shipping error, the prepreg was exposed to ambient conditions during shipping and partially cured. 

It was possible to fabricate composite panels and blade test articles with no apparent defects using this 

NASA/TM—2015-218340 3 
Figure 1.—Recommended cure cycle for Hexcel IM7/8551-7.

material, although it is predicted that material properties could have been reduced by the prior ambient 

exposure. As the purpose of this work was to evaluate relative performance of a baseline composite to 

composites with embedded material, it was decided that this material would be suitable for this work. 

A thermoplastic polyurethane veil was incorporated into the leading edge structure to evaluate its 

influence on the composite damage tolerance. The average areal weight of the TPU veil was 15 gsm and 

the average fiber diameter was on the nanometer scale. Interlayer veils used for toughening have been 

shown to improve composite fracture toughness (Ref. 8). 

Sheets of nickel-titanium shape memory alloy (95NiTi) were purchased from Johnson Matthey as a 

0.005-in. thick sheet. The thickness was chosen to approximate the cured ply thickness of IM7/8551-7. 

The martensitic to austenitic transition (actuation) temperature was approximately 100 °C, however this 

material was not preconditioned to change shape.

3.2 Test Article Geometry 

Material performance was evaluated through impact testing of a leading edge test article derived from 

the geometry of an aerodynamically optimized fan blade. The optimized blade utilized a modern fan blade 

design for a high-bypass ratio, low pressure ratio fan for a large commercial transport aircraft. This blade 

geometry was chosen to drive the design of the leading edge test article. The region represented by the 

test article is called out in Figure 2. The thickness profile of the test article was half scale of the thickness 

profile of the blade at 0.76 m (30 in.) fan radius, where the blade width was 30.5 cm (12 in.). This region 

of a blade was chosen as the FAA impact test requirements specify the impact location to be 50 percent or 

more from the root of the blade for large flocking bird impacts, where this profile chosen is 83 percent 

from the root (Ref. 1). 

The profile of the test blade is shown in Figure 3. It should be noted that the last 3.8 cm (1.5 in.) at the 

mid-blade location was adjusted to be flat and parallel in order to clamp into the test fixture. The profile 

was then extended in the blade radial direction (referred to later as the axial direction) to 45.7 cm (18 in.) 

in length, as shown in Figure 4, to create the leading edge blade test article. The profile was designed to 

compare various material modifications; not to demonstrate an optimum design in either geometry or layup configuration. The simplified geometry decreases manufacturing complexity while maintaining a 

leading edge profile. The changing thickness of the cross section requires ply drops across the blade from 

50 plies thick at the base to two plies at the leading edge. 

NASA/TM—2015-218340 4 
Figure 2.—Blade optimized for 

aerodynamics and highlighted cross 

section chosen for test article.

Figure 3.—Leading edge blade cross section with dimensions

Figure 4.—Leading edge blade shape and dimensions. 

NASA/TM—2015-218340 5 
3.3 Ply Lay-Up Configuration 

Laminate ply configurations presented in this paper were based on literature review (Refs. 5 and 6) 

and a symmetrical lay-up was chosen to minimize residual stress during the cure. The two goals of the 

blade ply configuration design approach included (1) high stiffness in the axial direction of the blade and 

(2) minimize design related impact damage to the leading edge. Three ply configurations and two ply 

drop patterns were used for the baseline design. 

3.3.1 Preliminary Ply Configurations 

Two separate lay-up designs were initially considered based on literature review (Ref. 5). These 

designs were labeled A1 and A2; as outlined in Table 1. 

The A1 and A2 design configurations were based on results reported in Reference 3; where leading 

edge fibers were oriented at a [0°/90°]s with respect to the axial direction of the blade. It was suspected 

that impact damage to this blade was minimal due to the increase in elastic modulus resulting from both 

0° and 90° ply orientations at the leading edge. While the configurations given in Table 1 describe the layup at the blade root, ply drops throughout the test article were chosen such that the leading edge ply 

configuration for designs A1 and A2 was [90°/–45°/0°/+45°]s with respect to the blade axial direction. 

The difference between A1 and A2 is the addition of six 90° plies at the center of the A2 thickness, a 

modification which was based on the results of A1 panel impact testing. In this study, initial panels were 

used to evaluate the impact test set-up and method, with a focus on impact induced damage rather than 

damage resulting from test configuration including boundary conditions. The ply orientations and 

locations of ply drops for design A2 can be seen in the cross section view, Figure 5; the length of each 

section is proportional to show the relative locations of the ply drops.

TABLE 1.—INITIAL PLY DESIGNS

Ply drop type Ply lay-up version Configuration Test articles 

A 1 [90°/(–45°/0°/+45°/0°)3/0°12]s C001-C002 

A 2 [90°/(–45°/0°/+45°/0)3/03/(90°/0°)3/0°3]s C003 

Figure 5.—The A2 ply lay-up, through thickness view. 

NASA/TM—2015-218340 6 
Both A1 and A2 ply configurations utilize 0° plies as the widest plies to prevent resin pockets from 

forming at the blade leading edge during cure. Placement of the 90° plies on the outer surface was chosen 

to increase chord-wise bending stiffness and limit bending deflections during impact. The outer 90° plies 

terminate on the outer surface, however this termination was later covered by leading edge protection and 

was not considered a risk for delamination on impact. Within the remaining blade thickness, the ± 45° 

angle plies were distributed between 0° and 90° ply orientation, thereby reducing the build-up of 

interlaminar stresses during cure. In design A2, 90° plies were added to help strengthen the root of the test 

article to prevent root failure due to chord-wise bending and through-thickness shear at the end of the 

clamped section. 

The gradient in thickness of the blade profile requires ply drops throughout the structure. In panels A1 

and A2, the ply drops were arranged in an accordion drop pattern; shown in Figure 6 for panel A2. This 

design was anticipated to reduce the shear plane that might develop, for example, with a step pattern ply 

drop. 

3.3.2 Secondary Ply Configurations 

Modifications to the initial blade ply configuration were made following impact test of the 

preliminary design. The lay-up configurations of A3 and B1 are provided in Table 2 with the difference 

being ply drop pattern and location of ply terminations. The accordion ply drop pattern was retained in 

panel A3, whereas B1 was fabricated with a defined ply drop pattern.

The B1 design was established to provide a systematic approach to the ply termination points, 

allowing better balance between +45° and –45° terminations across the ply drops. The only difference 

between designs A3 and B1 is that ‘A’ has ‘accordion pattern’ or random ply drops while ‘B’ is a 

systematic ply drop pattern. The ply drop location and the lay-up configuration of the B1 type is shown in 

Figure 6. 

TABLE 2.—SECONDARY PLY DESIGNS

Ply drop type Ply lay-up version Configuration Test articles 

A 3 [90°/(–45°/0°/+45°/0°)6/0°]s C005 

B 1 [90°/(–45°/0°/+45°/0°)6/0°]s C006-C008 

Figure 6.—The B1 ply lay-up design 50 plies total with symmetry.

NASA/TM—2015-218340 7 
3.4 Addition of Materials for Blade Aerodynamic Efficiency 

Two separate blades were fabricated to evaluate impact resistance of subcomponents containing 

(1) a thermoplastic polyurethane veil for interlaminar toughening, and (2) SMA strips for eventual camber 

change application. Six SMA strips were incorporated into the blade test article to evaluate their influence 

on impact damage tolerance. The strips were not ‘trained’ and therefore unable to change shape. Test 

specimen C007 included the embedded nickel-titanium strips at the locations identified in Figure 7. The 

5.1 cm (2 in.) long and 6.4 mm (0.25 in.) strips were placed in cut-outs of the 0° ply shown in Figure 6. 

The strips on the right half of the blade were covered by the polyurethane veil material on either side of 

the SMA. The veil was cut to extend beyond the SMA borders by approximately 2.5 mm (0.1 in.) around 

each side. 

The test article containing a thermoplastic polyurethane (TPU) veil interleave was prepared to 

evaluate impact resistance imparted by increased interlaminar strain capability of the structure. As seen in 

Figure 8, the toughening veil was 22.9 cm (9 in.) long and placed in the midsection of the length. Three 

layers of TPU were added to the blade in the locations labeled in Figure 9. Two of the layers were 7.6 cm 

(3 in.) wide and the other was 5.1 cm (2 in.) wide. The veil placement was chosen to cover a significant 

portion of ply terminations; reducing free edge stresses. 

Figure 7.—Location of SMA, right with polyurethane veil.

Figure 8.—Location of 7.6 cm (3 in.) polyurethane veil layer. 

NASA/TM—2015-218340 8 
Figure 9.—Polyurethane veil toughened specimen.

Figure 10.—Three-dimensional model of mated die mold. 

3.5 Fabrication

The IM7/8551-7 pre-impregnated tape was cut according to the ply lay-up designations for each 

panel. An aluminum mold was custom designed to fabricate the leading edge specimen described in this 

paper. The mold was a mated 7075 aluminum die, shown in Figure 10, and was fitted with a vacuum 

pump venting attachment and a channel to capture excess resin flow. All blade coupons were processed in 

an autoclave following the cure cycle recommended by Hexcel. The cured blades were machined on the 

45.7 cm (18 in.) sides for a smooth surface to secure in the impact test fixture. 

In an effort to protect the leading edge from extensive break-out damage, leading edge protection was 

applied in the form of an industrial aluminum foil adhesive tape. The aluminum foil tape was 0.25 mm 

(0.01 in.) thick and 5.1 cm (2 in.) wide and adds approximately 20 g of mass to the edge of the blade. The 

adhesive tape was placed 2.5 cm (1 in.) from the leading edge on the concave side of the blade, rolled over 

onto the convex side, then smoothed and flattened. A fully prepared test article is shown in Figure 11. 

NASA/TM—2015-218340 9 
Figure 11.—Prepared leading edge test article. 

3.6 Impact Test Method 

Impact testing was performed in the Ballistics Impact Laboratory at NASA Glenn Research Center to 

simulate the material response from bird strike at the leading edge. Impact testing was performed with a 

single-stage gas gun, consisting of a 7 m barrel and a 0.35 m2

 pressure vessel. The pressure vessel was 

loaded to a pressure of 1.2 MPa (175 psi), and the pressure was released using a burst disk. The projectile 

was housed in a cylindrical sabot for protection at the initial pressure release. Both were accelerated 

through the barrel by the release of pressure. The sabot was halted at the end of the barrel by a sabot 

arrestor and the projectile continued into the test specimen. High speed cameras were used to capture the 

impact of the projectile on the test article and determine the estimated speed. These tests were performed 

at speeds between 305 to 350 m/s (1000 to 1150 ft/s). A gelatin bird simulant was used as the projectile, 

containing ballistic grade gelatin and microspheres to approximate the density of a bird; ~ 0.9484 g/cm3

. 

The molded projectile was cylindrical, with a length of 7.6 cm (3 in.) and diameter of 3.2 cm (1.25 in.). 

The procedure for making the bird simulant was optimized for making approximately 600 cm3

 (11.7 in3) 

of gelatin and outlined in the Appendix. 

The projectile velocity and orientation of the leading edge test article were derived from computer 

simulation of the relative motion of a fan blade and bird during a bird strike event. The blade was 

clamped along the axial direction at the flat section for cantilevered impact testing. The leading edge 

specimen and fixture set up for impact testing is shown in Figure 12. The impact location was at the 

leading edge of the blade where the angle of impact was measured as 66° from the perpendicular to the 

projectile path at the thickest part of the blade. The blade tip was orientated approximately 24° off of the 

projectile’s path, shown in Figure 13(a), and schematically in Figure 13(b). Between the blade and the 

fixture there is a 15.2 cm (6 in.) section of clay added to reduce additional pressure loading on the fixed 

portion of the blade. This was added to channel the gelatin away from the clamped area. 

4.0 Results and Discussion 

Seven blades, labeled as C001 to C008, were impact tested with the gelatin bird simulant at a velocity 

of approximately 305 m/s (1000 ft/s). Test article C004 was omitted from this study as it contained a 

unique ply drop and configuration. The impact event led to a significant amount of damage at the leading 

edge of most test articles. The results of C001, C002, and C003 impact tests were used to optimize the 

test conditions and coupon ply configuration of subsequent test articles. 

NASA/TM—2015-218340 10 
Figure 12.—Impact testing fixture secured in impact chamber.

(a) (b)

Figure 13.—(a) Location of impact on the leading edge, and (b) Scaled schematic, top view, of the projectile path and 

the blade. 

NASA/TM—2015-218340 11 
The first test samples, C001 to C002, were fabricated with the A1 ply configuration 

[90°/(–45°/0°/+45°/0°)3/0°12]s. C001 received the most significant damage on impact—with some of the 

damage attributed to the test set-up. On impact testing, sample C001 broke where it was clamped in the 

fixture, as shown in Figure 14(a). The damage at the clamp was attributed to the gelatin projectile driving 

into the space between the blade and the securing clamp, increasing pressure at the clamp and resulting in 

failure. Additional damage included peeling of the aluminum tape leading edge protection at the impact site. 

As a result, a section of the leading edge was fully broken out. As a result of the leading edge protection 

failure, the 90° outer plies delaminated from the leading edge of the blade and were sheared off. 

Based on the results of the C001 impact test, the C002 set-up was modified to reduce pressure buildup at the clamp site by including a clay ramp to distribute the load of the projectile. The results of this 

minor change had a significant effect on the test results. C002 did not fail at the clamp, although a 

fracture line was apparent, and leading edge damage was limited to localized delamination at the impact 

site, Figure 14(b). The aluminum tape leading edge protection again peeled off on impact; however there 

was no global surface delamination of the outer 90° plies; as seen in the C001 test article. 

The third composite blade, C003, was fabricated with the A2 lay-up pattern, to decrease the 

possibility of shearing at the clamp. The A2 ply configuration ([90°/(–45°/0°/+45°/0°)3/03/(90°/0°)3/0°3]s) 

has added 90° plies at the root of the blade in an effort to strengthen this region and minimize damage at 

the clamp site. Fracture in the region of the clamp was again visible, and significant damage to the leading 

edge is shown in Figure 14(c). 

The C005 test article was fabricated with an A3 lay-up, [(0°/–45°/0°/+45°)6/0°]s, with accordion style 

‘A’ type ply drops. The general randomness of the ply drops resulted in an uneven mixture of +45°s and 

–45°s at ply drops. As a result, this test article had significant damage, including a broken out section at 

the leading edge measuring approximately 17.8 cm (7 in.) by 3.8 cm (1.5 in.), Figure 15. It was 

speculated that the leading edge section on the concave side of the blade failed at the outer major ply 

drop. The fracture location coincided with an uneven mixture of +45° and –45° plies; potentially 

compounding the damage. Following the initial impact, reflection of flexural waves traveling in the 

45.7 cm (18 in.) direction caused the leading edge material to peel away. 

(a) (b) (c)

Figure 14.—(a) Delamination and break out of C001, (b) leading edge of C002, and (c) broken out section 

of C003.

NASA/TM—2015-218340 12 
Figure 15.—C005 impact damage.

Figure 16.—C006, C007, and C008 after impact.

The ply configuration and ply drop pattern of articles C006 to C008 are described by the B1 lay-up; 

[(0°/–45°/0°/+45°)6/0°]s, and repeating ‘B’ pattern ply drops. Still, significant damage resulted on C006: a 

broken out section of 2.5 cm (1 in.) by 17.8 cm (7 in.) at the leading edge is shown in Figure 16. There 

was no global delamination in the blade only localized damage surrounding the broken out section and at 

the leading edge free corners. As in C005, the leading edge protection peeled off on testing. C006 served 

as a baseline compared to blades C007 and C008. 

NASA/TM—2015-218340 13 
Figure 17.—Close up of C007 impact zone,

impact location shown by red arrow, SMA 

strips with TPU indicated by black ovals, 

plain SMA strips indicated by white ovals.

Test blade C007 was fabricated to contain six strips of SMA ribbon. The ribbon was inserted into cutouts from the carbon fiber. Three strips were inserted just below the impact site, and three strips inserted 

just above the impact site (Figure 16, impact location shown with red arrow). For comparison, the strips 

below the impact side included a piece of the TPU veil (Figure 16, indicated by black ovals); as described 

in the experimental section. Leading edge damage was noted in this sample on impact; as shown in 

Figure 17. The section of broken-out composite measured 3.8 cm (1.5 in.) by 7.6 cm (3 in.) and visible 

propagation of damage was observed on the upper section of the blade where SMA had been inserted 

without TPU veil. The lower SMA pieces which were surrounded by veil. In this blade visible damage 

stopped near the second SMA ribbon. 

Test article C008 included the TPU veil interleave for damage tolerance. Three sheets of the 

interleave material were added across the major ply drops to improve interlaminar fracture toughness at 

these locations. Veil location was outlined in Figure 9. As seen from Figure 16, there was very little 

visual damage to the C008 test article. No delamination at the leading edge was noted in the interleave 

region. As with many of the test articles, a small delamination was observed near one end of the leading 

edge in an unmodified region due to flexural wave reflection. This failure mode was observed in many of 

the leading edge tests and is considered an artifact of boundary conditions, occurring after the initial 

impact. Figure 18(a) to Figure 18(c), which show thermography images used to characterize the post-test 

damage to the blade. 

NASA/TM—2015-218340 14 
The thermal image in Figure 18(a) shows damage to the baseline blade was limited to break-out at the 

impact site. The dark area along the leading edge is an artifact of the test, due to the thinness of the blade 

at the leading edge. Similar damage is seen within the SMA incorporated blade (Figure 18(b)), where 

impact energy was again absorbed at the impact site, leading to material loss. The image shows retention 

of material on the side of the impact closest to the SMA covered with toughening material. Additional 

work is necessary to assess the contribution (if any) of the thermoplastic veil to that material retention. In 

the toughened blade the leading edge maintained its integrity, allowing more energy to be transferred into 

the blade, resulting in greater overall deformation in the form of large oscillatory flexing. The increased 

bending of the blade in board of the impact location appeared to result in some delamination just 

beyond the placement of the toughening veil, as evidenced by the thermal image of the toughened 

blade (Figure 18(c)). Increased flexing and bending was observed in the toughened blade, relative to 

the baseline, in high speed videos of the experiments as shown in the embedded video, Figure 19. As a 

result, the mid-blade delamination appeared to only occur in the ‘toughened’ blade. 

(a)

Figure 18.—(a) Thermography images of the baseline composite blade.

Upper image corresponds to the back surface and the lower image 

corresponds to the front surface.

NASA/TM—2015-218340 15 
(b)

(c)

Figure 18.—Concluded. (b)Thermography images of the SMA incorporated 

blade. Upper image corresponds to the back surface and the lower image 

corresponds to the front surface. (c) Thermography images of the 

thermoplastic veil incorporated blade. Upper image corresponds to the back 

surface and the lower image corresponds to the front surface.

NASA/TM—2015-218340 16 
Figure 19.—High speed video of a composite leading edge test 

article impacted by a gelatin projectile impact; side view. 

5.0 Conclusions 

A simplified composite blade subcomponent was designed, fabricated, and impact tested to simulate a 

bird strike event. The influence of ply configuration and test set-up were evaluated such that impact data 

would reflect modifications to the test article ply structure and materials modifications. An interleave 

toughening approach was taken in an effort to reduce damage on impact and enable a reduction of 

composite blade thickness. The interleave approach lead to a significant reduction in damage on impact. 

In addition, interleave toughening materials showed evidence of damage mitigation in the SMA 

embedded coupon following impact testing. 

NASA/TM—2015-218340 17 

Appendix.—Synthesis of Gelatin Bird Simulant 

Prior to preparing the gelatin, remove the frozen molds to a room temperature location. Measure 

1200 g of deionized water and heat to 70 °C on a hot plate’s low heat setting with a stir bar. Measure 

300 g ballistic gelatin and slowly pour it into the heated water stirring and breaking up gelatin clumps 

making sure to remove the stirrer as little as possible. Once the gelatin mixture is fully dissolved and 

around 65 °C, remove from hot plate and place in a vacuum chamber to degas the mixture. The vacuum 

should be around 20 in-Hg without boiling the gelatin mixture for about 5 min. The vacuum required to 

boil the mixture is dependent upon the temperature before it is entered into the vacuum chamber. Once 

the mixture is clear, this is an indication that the mixture it degased and gasses have risen to the top of the 

mixture. Then, remove from the vacuum chamber and skim the top surface removing only the foam layer; 

remove the stir bar. Measure 57 g of phenolic microspheres. Using a dry whisk, gently break up large 

clumps. While stirring the gelatin mixture quickly on the hot plate, slowly stir in the microspheres as 

quickly as possible. Again, skim the foam off the mixture. Then pour it into the molds and refrigerate for 

solidification. 

NASA/TM—2015-218340 19 
References 

1. Federal Aviation Administration (08/07/09). Advisory Circular 33.76-1A, “Bird Ingestion 

Certification Standards.” 

http://www.faa.gov/documentLibrary/media/Advisory_Circular/AC%2033.76-1A.pdf Accessed 

August 5, 2014 

2. GE’s website: http://www.ge.com/in/aviation/GE90-commercial-aircraft-engines Access August 5, 

2014 

3. Friedrich, L. A. and Preston Jr, J. L., “Impact Resistance of Fiber Composite Blades Used in Aircraft 

Turbine Engines,” May 1973, NASA CR–134502, PWA-4727. http://ntrs.nasa.gov/ Accessed August 

5, 2014. 

4. Friedrich, L. A. “Impact Resistance of Hybrid Composite Fan Blade Materials,” May 1974, NASA 

CR–134712, PWA-TM-5022. http://ntrs.nasa.gov/ Accessed August 5, 2014. 

5. GE Aircraft Engine Group, “Impact Resistance of Composite Fan Blades,” December 1974, NASA 

CR–134707, R74AEG320. http://ntrs.nasa.gov/ Accessed August 5, 2014. 

6. Hexcel Product Data for HexPly 8551-7 Epoxy Matrix. 

http://www.hexcel.com/Resources/DataSheets/Prepreg-Data-Sheets/85517_us.pdf Accessed August 

5, 2014 

7. Dow, Marvin B., Smith and Donald L. “Properties of Tow Composite Materials Made of Toughened 

Epoxy Resin and High-Strain Graphite Fiber,” NASA TP–2826. July 1988. http://ntrs.nasa.gov/

Accessed August 5, 2014. 

8. Thomas K. Tsotsis, “Interlayer Toughening of Composite Materials,” Polymer Composites, 2009, 

pp. 70–86. 

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