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Integrated Computational Materials Engineering for Airframe Composite Structure Applications S. P. Engelstad, R.W. Koon, J.E. Action Lockheed Martin Aeronautics Company

J.M. Riga Lockheed Martin Advanced Technology Labs

A.M. Waas The University of Michigan

D. Robbins, R.W. Dalgarno Autodesk

A.R. Arafath, A. Poursartip Convergent Manufacturing Technologies, Inc. COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

1

Outline • Introduction • Model Definition – Convergent COMPRO – University of Michigan – Micromechanics – Autodesk ASCA

• Current Results – COMPRO Material Characterization – University of Michigan – Micromechanics Modeling of Curing ply strengths – ASCA Modeling of OHC/FHC, CSAI

COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

2

Introduction • Integrated Computational Materials Engineering (ICME) – AFRL and industry vision to reduce the material and process development cycle time and cost

• Integrated Computational Methods for Composite Materials (ICM2) is an AFRL, GE, and LM Aero composites ICME program – GE studying engine applications – LM Aero studying airframe applications

• LM Aero Airframe Goals – Tailoring composite damage tolerance and in-plane strengths to achieve maximum weight savings – Reduce material qualification costs and time spans • •

Today’s BMI systems took 20 years from the lab to its usage on F-22 The insertion of new materials technologies has become less and less frequent, and available materials are a constraint on the design


3

Introduction • BMI systems have experienced increased usage due to key structural design properties such as – Open hole compression (OHC) and compression-strength-after-impact (CSAI) – These key properties often “size” the acreage of the aircraft composite skins.

• The ICM2 program attempts to amplify the weight advantage of IM7/M65 BMI – Studying the autoclave cure cycle effects – Goal to optimize these critical design properties for aircraft weight

• GE/LM team has chosen – Cure Process Effects: Convergent Manufacturing Technologies’ COMPRO – Multi-scale progressive damage: Autodesk Simulation Composite Analysis (ASCA) software – Ply level strength effects of cure: University of Michigan (UofM) COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

4

Model Integration • The ICM2 program will utilize a digital framework to integrate the three modeling tools previously introduced • This framework is based on the model integration interface ModelCenter®, and will be built through a joint effort between GE and LM Aero • Goal is to show end-to-end integration from materials database through part structure performance to carry out process trade analysis • This presentation describes the individual components of the digital framework, as well as current results of the airframe portion of this program


5

COMPRO Analysis • COMPRO analysis will be performed to analyze the cure cycle of panel/part – Different cure cycles studied to reduce residual stress and optimize degree of cure – Thermo-mechanical properties and non-mechanical strains at the end of the curing process are passed to ASCA – Thermo-mechanical properties during the curing process are passed to UM code


6

Convergent’s COMPRO Process Analysis

COMPRO: a multi-physics composite processing plug-in for 3rd party FE solvers COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

7

UM-Micro Analysis • Micromechanics analysis will consist of progressive damage analysis at the unit cell level to determine resulting ply level strengths – Inputs • • •

From COMPRO: Temperature profile, degree of cure, and cure rate vs. time Constituent properties Resin strength test data as a function of cure cycle

– Outputs •

Ply level strengths as a function of cure cycle


8

UofM Micromechanics Based Modeling of Effects of Curing on Ply Strengths • Curing stress can cause micro damage that alters mechanical properties • Micromechanics analysis performed at the fiber-matrix scale to capture failure mechanisms – – –

3D hexagonally packed representative unit cell (RUC) and multi-fiber random packed unit cells RUC analyzed to compute the actual strengths of the lamina Optimize cure cycle to reduce detrimental effects of residual stresses

• UoM approach is based on two steps –

Step 1: • •

–

First the temperature profile, degree of cure and cure rate in the matrix are computed using the COMPRO software Determine stress evolution in the RUC due to the curing and the possibility of damage during cure

Hexagonally packed RUC

Step 2: • • •

Apply mechanical loading to the RUC in different directions to determine strengths Damage progression modeled using the crack band approach Outcome of this analysis is ply level composite stiffness and strengths, while accounting for the cure cycle


9

Autodesk ASCA Analysis • Autodesk Simulation Composite Analysis (ASCA) Progressive Damage analysis will be performed to determine the effect of cure on critical part strengths – GE will focus on engine duct flange strengths – LM will focus on Open-Hole-Compression (OHC), Filled-HoleCompression (FHC), and Compression-Strength-After-Impact (CSAI) allowable strengths Open Hole Compression

Delam 1

Ply 14: 0°

CSAI ASCA Model COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

10

Autodesk ASCA • ASCA software decomposes composite stress and strain states into constituent average stress and strain states – Stiffness of the damaged composite material is obtained by homogenization of the microstructure

• ASCA features two different methods for degrading the stiffness of damaged constituent materials – Instantaneous degradation of the stiffness of the damaged constituent to a user-specified fraction of the original stiffness – Energy-based degradation of the stiffness is gradually reduced to zero as the strain level increases beyond the damage initiation level


11

Current Results


12

Current Results • First year of ICM2 program Airframe efforts at the lamina level focused on – COMPRO material characterization of IM7/M65 lamina – UofM predictions of cure cycle effects on the lamina elastic properties and strengths – Physical resin “effect-of-cure-cycle” tests as required by UofM analysis

• At the laminate level, ASCA progressive damage tool being linked with COMPRO cure residual stress and UofM micromechanics strength property predictions – ASCA baseline cure cycle models developed for • • •

Open-Hole Compression (OHC) Filled-Hole Compression (FHC) Compression-Strength-After-Impact (CSAI)


13

Convergent Process Analysis Material Characterization • Mechanical constitutive model needed for process-induced residual stress development –

– – –

Simplest approach is “Cure Hardening Instantaneously Linear Elastic” (CHILE) model, where modulus of elasticity changes as a function of the instantaneous temperature and degree of cure (e.g. White and Hahn (1992) and Johnston et al. (2001)) But polymers show viscoelastic behavior, especially partially cured at high temperatures in a cure cycle CHILE models do not accurately capture the residual stress development during free standing post cure of a partially cured composite structure [Zobeiry 2003] Differential viscoelastic approach is being implemented in the upcoming next release of COMPRO (Version 3) to represent the viscoelastic behavior of a curing thermoset matrix composite

• An efficient methodology [Thorpe et al. 2013] to characterize the Hexply M65 material for viscoelastic constants was used here – –

Testing was performed on neat resin beams and uni-directional (UD) beams Note that all UD beams were prepared such that testing could be performed in the 2direction (in-plane, transverse)

Neat resin and unidirectional beam samples COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

14

Convergent Process Analysis Material Characterization • The COMPRO model was verified by comparing the spring-in angle of an L-shaped part manufactured on an invar tool using the baseline cure cycle • The model prediction of spring-in angle compared to the experiment is shown below

Model prediction of spring-in angle of an L-shape composite part compared to experimental results


15

Degree of Cure

• Process-induced residual stresses can be sufficiently high to cause damage in the material during the manufacturing process, or accelerate the formation and growth of cracks during the service conditions • While process-induced residual stresses are unavoidable, it was shown by [Li et al. 2014] that it is possible to reduce them by altering the cure cycle • In this work, evaluated this effect for Hexply M65 BMI laminates using the manufacture recommended cure cycle (MRCC)

Temperature, Tg

Convergent Process Analysis Effect of Cure Cycle Residual Stresses

Time

Manufacturer recommended onehold cure cycle for HexPly M65 simulated by RAVEN

– Convergent’s RAVEN process simulation software was used to predict the development of degree of cure (DoC) and the glass transition temperature (Tg) of the material for the MRCC COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

16

Degree of Cure

• As shown in other studies [Madhukar et al, 2000, White and Hahn, 1993 and Li et al., 2014], an intermediate hold can be added to the cure cycle to reduce the process-induced residual stresses • A three hold cure cycle was designed using RAVEN software such that the gelation and vitrification occur during the first hold • The second ramp rate was decreased to a small value to ensure that Tg always remains higher than the part temperature

Temperature, Tg

Convergent Process Analysis Effect of Cure Cycle Residual Stresses

Time

Modified three-hold cure cycle for HexPly M65 simulated by RAVEN


17

Convergent Process Analysis Effect of Cure Cycle Residual Stresses • An un-balanced cross ply [02/902] Hexply M65 laminate was analyzed for processinduced residual stress using the three-hold cure cycle

– – –

In both cycles, residual stresses are zero before gelation since resin modulus is negligible During isothermal hold, cure shrinkage coupled with the resin modulus development result in the development of residual stress Upon cooling down, residual stress further increases due to thermal shrinkage Process induced residual stress reduced by 6% for this preliminary alternate cycle Although a smaller value than shown by Li et al [2014], this reduction still significant compared to transverse failure strain of Hexply M65

Preliminary results of residual stresses in 900 layer of unbalanced cross-ply laminate along two different cure cycles

Residual Stress

– –

Temperaturue COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

18

UofM: Micromechanics Modeling of Effects of Curing on Ply Strengths UofM COMPRO model inputs to UofM Micromechanics Analysis

Temperature Profile

Degree of cure vs. time

from COMPRO

from COMPRO

Cure rate vs. time from COMPRO


19

UofM: Micromechanics Modeling of Effects of Curing on Ply Strengths • During curing, the matrix gradually solidifies, its stiffness increases, and the cell simultaneously contracts (cure shrinkage) due to network formation – –

Residual stresses develop in the matrix owing to cure shrinkage and thermal stresses Depending on level of tensile stresses developed, the degree of cure and the rate of cure, the material may crack locally during curing

• If damage occurs, its evolution is tracked until failure is reached • To date, no damage has been observed in this RUC analysis, by applying the cure cycle discussed • Subsequently, the “cured” RUC is subjected to mechanical load –

Example: Applying a tensile loading in 2-direction

RUC dimensions and mechanical loading applied


20

UofM: Micromechanics Modeling of Effects of Curing on Ply Strengths • After the peak strength, a significant reduction in stiffness occurs in RUC, and the crack path starts to be defined • Along the stress-strain curve the failed elements in the RUC are shown in red for four points (Points A, B, C, D) – – – –

Failure starts around the central fiber in Point A Elements start failing around the 2 fibers as shown in Point B Failure propagates as shown in Point C

Stress-Strain curve mechanical loading

Point A

Point B

Point C

Point D

At Point D the RUC is completely split along the red path

• This sequence of events has been verified for different finite element meshes and mesh objectivity is found to prevail


Crack path at Point A

21

ASCA OHC/FHC Baseline Cure Analysis Baseline Cure Cycle Preliminary Analysis

Aperture meshing for filled and open hole specimens COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

22

ASCA OHC/FHC Baseline Cure Analysis Baseline Cure cycle: Load-stroke plots comparing model and test data • OHC and FHC test panels for the T1 laminate [45/-45/45/-45/90/-45/0/45/-45/45/-45/45]s

T1 Laminate, FHC, Room Temperature

8 7 6 1.0 5 4

Test 1 Load

0.5 3

Test 2 Load

2

Test 3 Load

1

Model Load

00 -1

00

0.2 0.02

0.4 0.04

0.6 0.06

0.8 0.08

Actuator Displacement (in) Normalized Displacement

1.0 0.1

1.2 0.12

Normalized Compression Load Compression Load (kips)

Normalized Compression Compression Load (kips)Load

T1 Laminate, OHC, Room Temperature 1.5 12

10 1.0 8

6

Test 1 Load Test 2 Load Test 3 Load Model Load

4 0.5 2 00 00

0.2 0.02

0.4 0.04

0.6 0.06

0.8 0.08

1.0 0.1

1.2 0.12

1.4 0.14

Normalized Displacement Actuator Displacement (in)


23

ASCA OHC/FHC Baseline Cure Analysis Baseline Cure cycle: Load-stroke plots comparing model and test data • OHC and FHC test panels for the T2 laminate [45/90/-45/0]3s T2 Laminate, FHC, Room Temperature

1.5 12

10 1.0 8


6 0.5 4

2 00 00

0.2 0.02

0.4 0.04

0.6 0.06

0.8 0.08


1.0 0.1

1.2 0.12

Normalized Compression Compression Load (kips) Load

Normalized Compression Load Compression Load (kips)

T2 Laminate, OHC, Room Temperature 1.0 20

15 0.5 10


5 00

-5

00

0.2 0.02

0.4 0.04

0.6 0.06

0.8 0.08

1.0 0.1

1.2 0.12

Normalized Displacement Actuator Displacement (in)


24

ASCA OHC/FHC Baseline Cure Analysis

T3 Laminate, OHC, Room Temperature 20 1.0 15 0.5 10


5 00 00

0.2 0.02

0.4 0.04

0.6 0.06

0.8 0.08


1.0 0.1

1.2 0.12



Baseline Cure cycle: Load-stroke plots comparing model and test data • OHC and FHC test panels for the T3 laminate [45/0/-45/0/0/45/90/-45/0/0/45/0/-45/0]s T3 Laminate, FHC, Room Temperature 30 1.0 25 20 15 0.5


10 5 00 00

0.2 0.02

0.4 0.04

0.6 0.06

0.8 0.08

1.0 0.1

1.2 0.12

1.4 0.14



25

ASCA OHC/FHC Baseline Cure Analysis Baseline Cure cycle: Load-strain plots comparing model and test data • OHC and FHC test panels for the T2 laminate [45/90/-45/0]3s ed o e Co p ess o | [ 5/90/ 5/0]3s T2 Laminate, FHC, 75°FRoom Dry Temperature

75°F Room Dry Temperature T2 Laminate, OHC,

1.0 12



3s

18

1.0 16

10 8

0.5 6

Test 1 Load

4

Test 3 Load Model Load

00 00

1000 0.2

2000 0.4

3000 0.6

4000 0.8

5000 1.0

StrainStrain (µε) Normalized

6000 1.2

7000 1.4

12 10

Test 1 Load

0.5 8

Test 2 Load

2

14

8000 1.6

Test 2 Load

6

Test 3 Load

4

Model Load

2

00

0

0.2 2000

0.4 4000

0.6 6000

0.8 8000

1.0 10000

1.4 14000

1.2 12000

StrainStrain (µε) Normalized


26

ASCA OHC/FHC Baseline Cure Analysis • Baseline Cure Cycle: Comparison of ultimate load at failure between model and test, and the % error associated with the model values Laminate

OHC, % Error

FHC, % Error

T1

-8.0

-15.5

T2

5.7

-9.8

T3

2.3

-17.7


27

Compression Strength After Impact Model Baseline Cure Analysis • IM7/M65 Material • 55 ft-lb impact energy • Post impact NDI to determine damage present – Time of flight scan data shows the depth of the delamination created during impact

Impact Location

Outer Section (Single Layer) Inner Section (Multiple Layers)

Delaminations from Impact

• An ABAQUS model was built that matched what we could see in the NDI data – Multiple delaminations through the thickness – Virtual Crack Closure Technique (VCCT) to model delam growth during the analysis – ASCA for in-plane damage

Time of Flight Scan

Modeled

Inner Section (multiple sublaminates through the thickness) Delaminations from Impact

Impact Location Sublaminate 1 Sublaminate 2 Sublaminate 3 COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

Sublaminate 4

28

Compression Strength After Impact Test Data Overview • Compression strength after impact testing on impacted specimen Normalized Load

– 5x7 test fixture – Load and Strain response measured – Far field gages show a generally stable response – Divergence of near field gages indicate a damage propagation during the test

2/3

0.40 Test: SG 1 Test: SG 2 Test: SG 3 Test: SG 4

0.30 0.20 0.10 0.00

Specimen (green)

Normalized Load

4

Strain Gauges

0.20

0.40 0.60 Normalized Strain

0.80

Near-Field Strain Gauges

0.60 Test Fixture (grey)

5/6

0.50

0.00

Applied Load

1

Far-Field Strain Gauges

0.60

0.50

Divergence

0.40 0.30 0.20

Test: SG 5 Test: SG 6

0.10 0.00 0.00

0.20


0.40 0.60 Normalized Strain

0.80

1.00

29

Compression Strength After Impact Initial Correlation to Test Data Near-Field Back-to-Back 1.2 1.0

Normalized Load

• Initial stiffness correlation to test good • Ultimate load too high • Back-to-back predicted strains near the impact do not show divergence when the test does • Far field predicted strains show similar behavior as the near field predicted strains

0.8 0.6

Test: SG 6

0.4

Model: SG6

0.2

Test: SG 5 Model: SG5

0.0 0.00

0.50

CSAI Full Panel

1.50

2.00

Far-Field Back-to-Back

Fixture plate boundary condition (red)

1.2 1.0

1

5/6 2/3

4

Normalized Load

Test window

1.00 Normalized Strain

0.8 0.6

Test: SG 2

0.4

Model: SG2 Test: SG 3

0.2

Model: SG3

0.0 0.00

0.50


1.00 Normalized Strain

1.50

2.00 30

CSAI Model Next Steps • Need to greatly improve fidelity of NDI and test instrumentation • Collaborative effort with AFRL/RX – Synergy with existing internal RX program AFRL/RX perform detailed Computed Tomography (CT) Scans • Goal to very carefully measure post-impact state • AFRL/RX compression test three specimens and use Digital Image Correlation (DIC) on both sides of specimen to identify displacement and strain field for improved model correlation •

– AFRL/RX to model specimens in BSAM, ICM2 to model in ABAQUS, and compare results


31

Compression Strength After Impact • Panel Level Testing will be performed to Validate the Effectiveness of Cure Directed Modeling Efforts Test Specification

Layup

# of Plies

Specimen Size

Short Beam Shear

D2344

[0]50

50

1.5"x0.5"

Double Cantilever Beam

D5528

[0]26

26

9.0"x1.0"

3

3

3

90°/0° Compression

D6641

[90/0]7s

28

5.5"x0.5"

3

3

3

In-Plane Shear

D3518

[45/-45]2s

8

9.0"x1.0"

3

3

3

Open Hole Tension

D5766

[45/90/-45/0]3s

24

12.0"x1.5"

3

3

3

D5766/D6742 [45/90/-45/0]3s

24

12.0"x1.5"

3

3

3

[45/90/-45/0]3s

24

12.0"x1.5"

3

3

3

D6484/D6742 [45/90/-45/0]3s LMA-PT001 Compression Str. After Impact [45/90/-45/0]4s Method 4.30

24

12.0"x1.5"

3

3

3

32

11.0"x13.0" Impact 10.0"x12.0" Comp.

3

3

3

Laminate

Lamina

Test

Filled Hole Tension Open Hole Compression Filled Hole Compression

D6484


Test Conditions -65°F Dry 70°F Dry 350°F Wet* 3 3 3

32

Summary • Results of the airframe portion of the GE/LM Aero ICM2 program – Studying the IM7/M65 bismaleimide (BMI) system for application to next generation airframes

• Linking general capabilities of three analysis tools – Convergent COMPRO – Autodesk ASCA – University of Michigan Micromechanics

• Current results include – COMPRO analysis of two cure cycles for the IM7/M65 system – University of Michigan micromechanics analysis of failure strengths in an RUC for this system – ASCA progressive damage analysis of OHC, FHC, and CSAI specimens

• Integration of these codes will enable: – Laminate level damage analysis including effects of varying cure cycles COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

33

Convergent Process Analysis Material Characterization • Testing was performed using a TA Instruments Q800 Dynamic Mechanical Analyzer (DMA) using 3-point bend geometry with a 50mm span – – –

During each test, the complex moduli (storage modulus, loss modulus and tan delta) were measured Two types of tests were performed; constant frequency temperature ramps and isothermal frequency sweeps. The constant frequency temperature ramps examine the temperature dependent moduli and the Iso-thermal frequency sweep tests examine the frequency dependent moduli

Storage Modulus

• A generalized Maxwell model was used to predict the viscoelastic modulus of HexPly IM7/M65 Material

Temperature

Prediction of storage modulus using the developed Maxwell model COPYRIGHT 2014 LOCKHEED MARTIN CORPORATION – ALL RIGHTS RESERVED

35

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