Lessons Learned in Full Composite

Aero-Structures Development in Japan

ICAS “Lawrence Hargrave” Lecture

for Innovation in Aeronautics, Sept. 25/ 2012

Professor, Aerospace Engineering,

Graduate School, Nagoya University

(Formerly Executive Director,

Japan Aerospace Exploration Agency)

Dr. Takashi Ishikawa

Outline of the Lecture

• Foundation of Mechanics of 2-D Textile Composites: Leading the World(1980-1985) Started at NAL, Continued in USA Stay (University of Delaware with Prof. Chou)

– Motivation: Full Composite Tail Plane Model of NAL STOL Research Aircraft

8 Harness Satin was Employed. Serious Warping in On-Ply Consolidation– Key: ng in Weave, Mosaic Model, Crimp Model, Bridging Model

・ Buckling and Post Buckling Compressive Failure of T-Stiffener

– Motivation: Immature Failure of Full Composite Tail, Buckling Induced Failure

– Nonlinear Analysis of T-Stiffener: Good Agreement with Experimental Results in CF/Epoxy

• Challenge to Clarification of Mechanics of Compression after Impact (CAI)

– Observation of Delamination Propagation, Clarification of Final Failure Process

– Numerical Simulation with Partner, Professor Suemasu: Cohesive Zone Element

• Mechanics of Compression after Impact (CAI) in Stiffened Panels

– Good Correlation of Numerical Prediction with Experimental Results

• Clarification of Interlaminar Strengthening Mechanism by Stitching

– Conception of Single Stitch Pull-Out Test and Observation by Micro X-Ray CT

• Contribution to Strength of Japan’s Aero-Industries in Composite Technology

• A Typical Example: 35% Work Share in Boeing 787 Production, Mostly Composites Portion, Particularly Main Wing

1st Topic: Mechanics of 2-D Textile

Composites, Triggered by Full CFRP

Horizontal Tail for NAL STOL R&D Aircraft

Initial Development: Sinusoidal Web Spar

⇒⇒⇒⇒8 Harness Satin Employed for Better

Drapability. Bending Failure in Web

Final Development: Model of Full CFRP

Horizontal Tail ⇒⇒⇒⇒8 Harness Satin, Used

in Front and Rear Spar

Extensive warping was generated by 1 ply 8-H Satin. ⇒⇒⇒⇒

The reason was unknown at that time: Request of consulting by

manufacturer ⇒⇒⇒⇒ Cause was easily estimated by observation of

sectional view of 8H Satin: Gate to “Treasure Island”

Pattern of 2-D Fabrics: Relationships between

Interlaced Region and Geometrical Repeat Number = ng

(a) ng=2: Plain Weave: Basic/No Anti-Symmetricity (b) ng=3: 2/1Twill Weave

(c) ng=4: 4 Harness Satin (Claw Foot Satin) (d) ng=8: 8 Harness Satin

Idealization Process to the Mosaic Model

Left: (a) Fabric Only, (b) After Consolidation, (c) Neglecting Crimp

Right: (a) Mosaic Model for 8H Satin, (b) One Base Unit, (c) Alignment

Transverse to Load, (d) Alignment Parallel to Load

Upper / Lower Bounds and Solution by Crimp

Model for Coupling Compliance b*11

Coupling Compliance

vanishes at 1/ng= 0.5

(Plain Weave)

○○○○,●●●● correspond to 2-

D FEM results by

Parallel Mosaic Model

and Crimp Model,

respectively.

Results by Crimp Model

and Parallel Mosaic

Model almost coincide

with each other.

Crimp Model Concept for Mechanical

Properties of Fabric Composites

Accomplishment at

University of Delaware

Lenticular Section

Shape of Warp (y-

direction) is assumed

Neat Matrix Region is

assumed ⇒⇒⇒⇒Slight

Difference in Whole Vf

and Thread Vf

Sinusoidal Crimp

Shape is assumed

Upper and Lower Bounds of In-plane Stiffness A11

and A11 Results by Crimp Model

Upper and Lower Bounds

are Most Separated at

1/ng= 0.5

○○○○,●●●● Corresponds to

2-D (longitudinal and

thickness directions) FEM

Results of A11 for Serial

Mosaic Model and Crimp

Model, Respectively

One Dimensional Crimp

Model Provides Quite Low

A11 Value

Bridging Model for 8 Harness Satin

Basic Idea of the Model

Plain Weave Element Showing

Low In-plane Stiffness is

Surrounded by Fiber Cross-Ply

Element with Straight Fiber

Hexagonal Shape of Actual Unit

Cell of 8H Satin is Replaced by

Square with the Same Area

(This Model is only Feasible at

ng≥4)

Regions B, C, and D are

Assumed to Share the Same

Reference Strain and Curvature

Comparison of A11 by Bridging Model

and Crimp Model

Bridging Model Results Fall between Upper and Lower Bounds

Bridging Model Results Agree with 3-D FEM Results for 8H Satin

Comparison between Experiments and Theoretical

Predictions of Elastic Moduli of Textile Composites

Experimental Results were

Obtained in NAL

Dependency of Elastic Modulus

on Ply Numbers in Plain Weave:

Effect of Restraint of Out-of-

Plane Deformation by Stacking

of Multiple Layers

Experimental Results of 8H

Satin: Obtained by Back-to

Back 2 Ply Laminates

Predictions Coincides Basically

with Experiments

By-Product of Non-Linear Behavior Research of Textile

Composites in Tension: Longitudinal Hardening Non-

Linear Behavior of UD CFRP

Tensile Non-linear Behavior of Plain Weave Composites in Strip (one thread) Specimen

Cannot be Explained by Geometrical Non-Linearity Only ⇒⇒⇒⇒

Hardening Nonlinearity of CFRP in Fiber Direction must Be Introduced

[0]8ply Stress-Strain Curve

0

500

1000

1500

2000

2500

3000

0 2000 4000 6000 8000 10000 12000 14000 16000 18000

Strain (µε)

Str

ess

(M

Pa

)

IM600/Q -C133

Constitutive Equation Based on Higher Order Elasticity: Expression of Elastic

Modulus by Fractional Function of Stress of the Second Order

EL = 1/ (S11 + 2S111 σ1 + 3S1111 σ12) [Ishikawa, et al.: J. Mat.Sci. Vol.20]

UD-CFRP using Toray T-300: EL = 1000/ {6.689 + 0.982(σ1 – 1.10)2} [GPa]

By-Product of Non-Linear Behavior Research of

Textile Composites in Tension: Longitudinal

Hardening Non-Linear Behavior of UD CFRP

----1.01.01.01.0 0000 1.01.01.01.0

160160160160

140140140140

120120120120

100100100100

Equation 18

Equation A5

From[18]From[18]From[18]From[18]

Present workPresent workPresent workPresent work

Stress (GPa)Stress (GPa)Stress (GPa)Stress (GPa)CompressiveCompressiveCompressiveCompressive TensileTensileTensileTensile

（Lo

ngitu

dina

l Ten

sile

Mod

ulus

Long

itudi

nal T

ensi

le M

odul

usLo

ngitu

dina

l Ten

sile

Mod

ulus

Long

itudi

nal T

ensi

le M

odul

us）） ））

(GPa

)(G

Pa)

(GPa

)(G

Pa)

Non-linear Elastic Behavior of UD-CFRP

(Using Toray T-300) : Equation Explained

EL = 1/ (S11 + 2S111 σ1 + 3S1111 σ12)

Recent Paper with Yokozeki: Compressive

Behavior by Modified Sandwich Beam

Longitudinal Non-Linear Behavior

#29

Serious Elastic Modulus Reduction

at High Compression Strain

Longitudinal Hardening Non-Linear Behavior of

UD CFRP

Clarification of Compressive Failure Behavior

of Stiffeners of Aircraft Wing: Motivated by

Immature Failure of Full CFRP Horizontal Tail

After Bending Test at Ultimate Load by

Symmetric Down Forces at Wing Tips:Failure Location was Lower Flange in Aft Spar

Close-up Photo of Aft Lower Flange Spar

⇒⇒⇒⇒ Typical Buckling Induced Failure

Immature Failure at 74% of Predicted

Load ⇒⇒⇒⇒ Insufficient Understanding of

Spar Flange Buckling in Corrugated

Web Spar in Sine-Wave ShapeStress Strain Data near Failure Point

One More Incentive of Buckling Induced Failure,

Failure in Spar of “HOPE” Reentry Vehicle

Spar Material : CF/Polyimide, Failure was Induced by Buckling

Due to Inaccuracy of Temperature Prediction, too much Safety Factor ⇒⇒⇒⇒

Heavy Structure, Requirement of Precise Prediction of Buckling

Employed Specimen Geometry: T-Stiffener

Photo of Typical Buckling Deformation

Specimen : T-Stiffener = Element of Spar Flange or Stiffened Panel

Used Material : CF/Epoxy and CF/PEEK (Thermoplastic Composites)

Sectional Shape of Specimen and

Important Dimensions

So Called Unidirectional “Corner Filler: UD Noodle”””” is Inserted

If Specimen is Thin, Difference between bw and b’w = bw – t/2 is negligible

If Specimen is thick, This Difference should be Noticed

Modeling of Intersection Area is One Key Issue

Approximate Solution by Rayleigh-Ritz Method

Points: Sinusoidal in x (Longitudinal), Sine Hyperbolic in y (Transverse)

Purpose: Explicit Survey of Effect of D16,D26 (Bending- Torsion Coupling

Terms)

22

66

2

0 2

2

222

2

2

2

12

2

2

2

11 422

1

∂∂

∂+

∂

∂+

∂

∂

∂

∂+

∂

∂= ∫ ∫ yx

wD

y

wD

y

w

x

wD

x

wDΠ

a

o

b

p

dydxx

wN

yx

w

y

wD

yx

w

x

wD x

∂

∂+

∂∂

∂

∂

∂+

∂∂

∂

∂

∂+

22

2

2

26

2

2

2

16 44

Potential Energy under Out-of-Plane Deformation w, Πp, can be written as follows

Out-of-Plane Deformation Function, w(x,y) of One Piece of T-Stiffener,

Rectangular Strip of Three Simply Supported Edges and One Free Edge is

Assumed in the Following

+

⋅

=∑∑∑

= = = b

ypB

b

nyA

a

xmyxw mpmn

m n p

ππsinsinhsin),(

3

1

3

1

3

1

Influence of Thickness Disturbance on Buckling

Prediction of T-Stiffener: Finite Element Analysis

Used Mesh and Predicted

Buckling Mode:

Close to Experiment

In Case of CF-PEEK, Serious Thickness

Scatter Took Place due to Immature

Process Technology (Late 1980)

Predicted Buckling Deformation in Web( for W Type)

Considered Important Matters: Geometrical Nonlinearity (Analysis for

Deformed State), Appropriate Initial Imperfection, Precise Boundary

Conditions along Loading Edges, Mode of Loading

Above Results: Iso-value Contour of Out-of-Plane Deformation in Web

240.8MPa 356.0MPa 386.8MPa 416.6MPa 446.3MPa 475.6MPa

Comparison of Predicted Nonlinear Strain with

Experimental Results (for W Type)

Predicted Strain at Exact

Gage Location : 5mm Inside

from Free Edge at Central

Point in Terms of Loading

Direction

Loading Edge Boundary

Condition: Fixed is

Appropriate

Very Good Agreement with

Experimental Results

Predicted Failure Strength of T-Stiffener and

Experimental Results for CF/EpoxyPredictable if Consecutive

Failure does not Happen

Ratio of Width to

Thickness :

[(bf-t)/2]/t = 6.5 is Optimal,

Threshold between

Buckling Failure and

Material Failure

If This ratio is Larger than

6.5, Early Buckling

induces Lower Strength

Critical Ratio Depends on

Material

Buckling Critical Load of Sinusoidal Web Spar

Buckling is Sensitive to Distance to Free Edge

Numerical Results of Buckling Mode

Buckling Takes Place at Bottom of Sine

Length of Spar is

Different from Tested

Model

Prediction in 1982 was Done

Based on Flat Web due to

Computational Resources

Clarification of Mechanism of

Compression after Impact (CAI): Concept of CAI Test

Impact at Required Energy Level ⇒⇒⇒⇒ Determination of Delamination Area by

Ultrasonic NDE ⇒⇒⇒⇒ Compressive Strength Measurement

(a) Impact by Drop Weight (b) Non-Destructive Evaluation (c) Compression

Impactor

SpecimenDelamination

Compressive

Force

Strong Incentive of CAI Research:

7J7 Horizontal Tail Test Program

From 1988 to 1991, as Boeing-JADC-NAL(at that time) Cooperative Research

Test Venue: NAL Chofu , under Witness of Boeing

7J7: Code Name

of Imaginary

Aircraft

Actually

Prototype of

Boeing 777 Full

Composite Tail

Standard

Development

Sequence was

Taken

Outline of Full Composite Tail of 7J7,

Designed by Boeing,

Fabricated by Fuji Heavy Industries

Rough Drawing of the

Model and Basic

Dimensions

Photo of Robotic Ultrasonic C-Scan of Six

Impact Locations before Fatigue Loading

(Maximum Energy: 900lbs*in)

and Trigger Point of the Final failure and

Skin Split Line

Trigger Point of

the Final Failure

Split Line

Acoustic Emission Data of the

Final Failure Test of 7J7Full Composite Tail

The Final Failure was Triggered by Impact Point 3B (850lbs*in) [Most Active AE Signal

Generation]. Compression Failure with Delamination Propagated Instantaneously to

the Transverse Direction to the Main Load.

Wing Skin was Split like by Axe Strike (No Picture by Boeing Control)

Large Scale CAI Failure Happened. Importance of CAI was Strongly Believed

Essence of CAI Research Results:

Typical Delamination Damage Geometry

Impact Damage of Less Tough CF/ Epoxy

at High Energy Level (by NASA Method of CAI)

Example of Hat and Brim Shape Delamination

0

1000

2000

3000

4000

5000

0 1 2 3 4 5 6 7

SACMA STD

SACMA PF

NASA

Normalized Impact Energy (J/mm)

Del

amin

atio

n A

rea

(Pro

ject

ion

, mm

*2)

Essence of CAI Research

Relationship between Impact Energy per Thickness

and Delamination Area (Projection)

Energy Threshold for Delamination Creation is Identified

Leveling-off in Delamination Area at High Energy:

Delamination Edge is Approaching to Supporting Fixture

Relationships between Compression Stress

and Out-of-Plane Deformation at Delamination Center

in CAI Test ( CF/Epoxy & CF/PEEK)

All Delamination

Pieces Deformed

to the Same

Direction: Mode II

Component is

Dominant

Deflection Reverse

at Delamination

Pieces at Impact

Side : Mode I

Component is

Dominant

Delamination Buckling Behavior

and Measured Deformation

Consecutive Moire-Topography Pictures during CAI Tests

Delamination Propagation Transverse to Load

Both NASA

Method

Upper:

CF/Epoxy

Lower:

CF/PEEK

Percentage:

To failure Load

0

100

200

300

400

500

0 2 4 6 8

SACMA STD

SACMA PF

NASA

Normalized Impact Energy (J/mm)

CA

I Str

eng

th (

MP

a)

CAI Properties: Relationships

between Impact Energy per Thickness and

Compression Strength Reduction

Different Tendency in NASA and SACMA methods

Incentive to Find the Other Interpretation Parameter

Another Interpretation Parameter: Summation of

Delamination Area by Using B-Scope of Ultrasonic C-Scan

Geometrical Relationships about Imaginary Cylinder

Diameter of Accumulated Delamination

Rc = Ac/As = tDc /(t b) = Dc / b, Re = Ae/As = tDe /(t b) = De / b

0

100

200

300

400

500

0 0.1 0.2 0.3 0.4 0.5

SACMA

NASA

Sectional Area Ratio of Virtual Delam. Cylinder: Rc=Dc/b

CA

I Str

eng

th (

MP

a)

Relationships between Ratio of Imaginary Cylinder

Area of Delamination to Specimen Width and CAI Strength

Dc/b as Interpretation Parameter: Experimental CAI Strengths of

NASA and SACMA Methods Fall around the Same Curve

H. Suemasu, W. Sasaki, T. Ishikawa, Y. Aoki, A numerical study on compressive

behavior of composite plates with multiple circular delaminations considering

delamination propagation, Composite Science and Technology, 68, 2008, 2562-2567

Relationships between the applied load and the deflections of a delaminated portion (B) and

undelaminated portions (A and C) for a plate with four delaminations(4D40).

Assumed Circular Multiple

(Four) Delamination

Important Nature of the

Solution: Captured Load Drop

with Delamination

Propagation

By Suemeasu, Ishikawa et al.

Numerical Results Based on

Multiple Delamination with

Professor Hiroshi Suemasu,

Sophia University, Japan

Delamination Propagation Analysis Using

Cohesive Zone Element

Interface 1

Interface 3

Interface 2

Interface 4

59.71 kN 57.55 kN 52.93 kN60.50 kN 61.57 kN 60.61 kN 60.56 kN

40 mm

Loading Direction

Numerical Results Based on Multiple Delamination with

Professor Hiroshi Suemasu, Sophia University, Japan

H. Suemasu, W. Sasaki, T. Ishikawa, Y. Aoki, A numerical study on compressive

behavior of composite plates with multiple circular delaminations considering

delamination propagation, Composite Science and Technology, 68, 2008, 2562-2567

By Suemeasu, Ishikawa et al.

Important Nature

of the Solution:

Transverse

Propagation of

Delamination to

Loading Direction

Experimental Fact

without Exception

Challenge to Clarification of CAI Properties

in Stiffened Panel: Initial Target = Buckling Analysis

Results of CF/Epoxy Stiffened Panel for Reference = Non-Damaged Part of STOL Tail Wing

Final Purpose = Verification of High Damage Tolerance of CF/PEEK Structures

Comparison of Numerical and Experimental

1st Buckling Mode

Good Correlation was Obtained for Both

Buckling Load and Mode by Including

Aluminum End Platen Potted into Analysis

Photo for CF/PEEK with 4 T-Stiffeners

Relationship between Impact Energy per Thickness

and Delamination Area in Stiffened Panels

In CF/Epoxy Panels (Low Fracture Toughness), Delamination Area Tends to Narrower

over Threshold: Penetration Mode of Projectile is Dominant

CF/PEEK (High Fracture Toughness): Still in Proportional Region (under Threshold),

Rather Wider Delamination Area Complicated Findings

200

220

240

260

280

300

0 50 100 150 200 250

FEA (Clamped)FEA (Simply Supported)Experiment

Temperature (Deg. C)

Final Goal Change: Main Wing of Supersonic Research Aircraft (1996-2003): Adopted

Material = CF/PIXA (Heat Resistant Thermoplastic) from Stiffened Panel to Wing Model

Photo: High Temperature Compression of Panel,

Comparison of Numerical Results and Experiments

Material Change to CF/PIXA (Heat Resistant

Thermoplastic): CAI Properties at High TemperatureDesign Change of Specimens: Awareness of

Cost and Improvement in Buckling Mode = 2 T

Stiffeners and End Support Fixture

0 2 4 6 8 10

0

100

200

300

400

500

Normalized Impact Energy (J/mm)

CAI S

treng

th (MPa)

CF/Epoxy,4T

CF/PEEK,4T,FY89

CF/PIXA,2T

Concluding Experimental Results:

CAI Properties of Stiffened Panels (at Room Temperature)

Compression over

200ﾟ゚゚゚ C Epoxy Potting:

Not Available

End Fixture was

Designed

Dimensions of Panel:

Inter-Bay Width and

Length, Improved to

Realize 4 Half Waves

at the 1st Mode

Conclusion of CAI Properties of

Composite Stiffened Panels

Overwhelmingly Excellent Results of

CF/PIXA, Almost Insensitive to Impact

Vanished due to Cost, Unfortunate

Research of Mechanics of

Interlaminar Strengthening by Stitching

Concept of Stitching and basic

Geometrical Parameter

Photo of Consolidated Flat Panel

after Stitching

(Carbon Fiber Stitching)

(1983-2007)

Stitched by Mainly Kevlar® Threads,

Some are by Carbon Threads with

Special Technology

4.0（Ref.）

0.30

A-AStitch Thread

Interlocking point

A A

□12.5□5.0(Ref.)

needle

bobbin

Conception of Single Stitch Thread Pull-Out

Specimen and Observation of Behavior by

Micro X-ray CT

With Stitch

Thread

0

100

200

300

400

500

600

700

0.00 0.50 1.00 1.50

開口変位（mm）

荷重

（N

)

Interlaminar Failure of CFRP

Failure of Stitch Thread

Pull-Out of Thread from CFRP

Load Carried by CFRP

Stitch Thread

Tensile Deflection (mm)

Load

(N)

0.0

2.0

4.0

6.0

8.0

10.0

12.0

14.0

16.0

0.0 1.0 2.0 3.0 4.0

Vst (%)

GIR

(N

/m

m)

Experimental Withput pull out thread effect With pull out thread effect

GIR-Vst plot with DCB test data

Relationship between Mode I Energy Release Rate and

Volume Fraction of Stitch Thread:

Comparison of FEM Simulation with Experiments

History of Composites Application to Aircraft

Cost Limits in

Commercial

Applications

If stick to

Traditional

Airbus A380

Boeing

B 787

Material Usage in Boeing B-787

Carbon laminate

Carbon sandwich

Fiberglass

Aluminum

Aluminum/steel/titanium pylons

Building on Proven MaterialsBuilding on Proven Materials

Composites

50%

Titanium

15%

Other

5%Steel

10%

Aluminum

20%

Courtesy of Boeing

←Mitsubishi Heavy I ↑Fuji Heavy I

Photo of B787 Main Wing Production

Contribution of the Award Recipient

Affects Somewhat This Japan’s

Production Share

Japan’s Aero-Industries Obtained 35 %

Production Share in Boeing B-787: Mostly Composites

Impressive Photo of Ultimate Test of

B-787 Full Composite Main Wing

Courtesy of Boeing Website

Lessons Learned in Full Composite

Aero-structure Development• Foundation of Mechanics of 2-D Textile Composites: High Impact Factor Papers

– One Drawback: Interruption after Clarification of Fundamental Mechanics• A Honor of Completion of the Theory Went to Dr. Naik, India

• Buckling and Post Buckling Induced Failure of Stiffener– Nonlinear Analysis of Postbuckling Behavior: Simple Maximum Stress Theory

Provides Prediction of Compressive Strengths = Good Agreement with Test– Finding the Optimal Width/Thickness Ratio in Stiffener

• Clarification of Compression after Impact (CAI) Mechanics– Introduction of High Performance Ultrasonic C-Scanner and Moire-Topography

Camera played a Key Role = Observation Equipments are Important

– Theoretical Predictions: Assistance by Professor Suemasu, Leading World• Mechanics of Compression after Impact (CAI) of Stiffened Panels

– Double Incentive: Demonstration of Damage Tolerance of Thermoplastic Composites, CF/PIXA Shows Extremely Excellent Insensitivity to Damage

• Cost (Material and Tooling) is Serious Issue = Forced to Quit

• Mechanism of Interlaminar Strengthening by Stitching– Pull-Out Tests of Single Stitch Thread: Understanding of Energy Dissipation

• In Conclusion, Accumulation of Lessons Learned by Award Recipient Raise Potential of Japan’s Aero-Structure Industries in Composites Area

• A Typical Proof: 35% Work Share in Boeing 787 Production by Composites

51

Thank you! for your kind attention