Aerospace Composites

Ramesh Sundaram

Head, Advanced Composites Division

National Aerospace Laboratories

Bangalore, India

CSIR-NAL

60’s

70’s

80’s

90’s2000-

The Beginnings

The first Blowdown

Strategic Missions

India’s First Super Computer

KodihalliBelur

HANSA

SARAS

CSIR-NAL

Turbo machinery

and combustion

Structural design,

analysis and testing

Aeroelasticity

Computational & Experimental AerodynamicsAerospace

electronics and systems

Flight mechanics & controls

Electromagnetics

Parallel Computing

Composites

Aerospace materials

Core competence of NAL spans practically the whole aerospace sector

What are composites

Why Use Composites ?

Composites have higher strength than traditional materials….due to aligned fibers carrying the load

Why Use Composites ?

Composites are stiffer than conventional materialsof the same weight due to their adoptive nature onecan align fibers in the direction to carry the load

Why Use Composites ?

Composites are lighter than traditional materialsdue to their tailorability they can be designed to minimum weight without sacrificing strength

Weight??

• Lower weight implies more freight

• 1 kg saved is equivalent to thousands of dollars over the lifetime of the aircraft

• In space craft 1kg of weight implies saving of 2 tonnes of propellant from earth to MARS

Reduce weight by using better materials and efficient design concepts

REQUIREMENTS OF STRUCTURAL MATERIALS

• Density as Low As Possible• High Strength & Stiffness• Should Have Better Corrosive Resistance• Should Not Be Fatigue Sensitive• Should Have Good Damping Properties• Should Have Good Damage Tolerance• Capable of Giving Good Surface Finish• Should Have Shape Retention Capability• Should Cost Less• Should Be Possible To Repair• Adapability To New Technologies

OTHER ADVANTAGES • Better Material Utilization• Lower Tooling And Maintaining Cost• Easier Parts Integration• Inhibited Flaw Propagation• Reduced Maintenance Cost• Low Energy Requirement For Production Of

Materials And Parts• High Corrosion Resistance• Antimagnetic And Low Radar Cross-section• Low Thermal Expansion• Excellent Surface Quality• Not Sensitive To Fatigue Loads• Excellent Shape Retention Capability

• No single material can provide all this

• Need to tailor the material by combining different materials

• A composite material is a physical combination of different materials which can perform in concert by taking the desired shape specially tailored to meet the requirements.

Whenever human beings encounter problems they will look into nature for solutions

• Interestingly nature has done these things long back

• Nature is a super designer, technologist, teacher and guide

Nature- The Super Designer

• Nature has realised long back to make efficient structures it has to go for composite structures

Composites can be broadly classified into two groups:

• Fibre Reinforced Composites

– Wood is a fibre reinforced composite -Cellucose fibres and lignin matrix

– The joints between branches and the trunk is superbly done by interlacing

• Particle reinforced composites

Another example is the Bone - Hybrid Composite

• It contains fibre reinforcement

• It has particle reinforcement

• Bone consists of organic fibres called Collogen and an inorganic matrix hydroxyapatite

The material is designed so as to have different properties in different locations.

• Inner Rings are softer than outer rings

– It can take care both dynamic and static loads

– Unlike the tree, the bone experiences a very complex three dimensional loading

• Teeth- Ceramic Composite made at room temperature !!!

Early Innovations by Humans

• Bricks of Clay reinforced with Straw ( Egypt -800 BC)

• Plant fibres worked into pottery

• Mummy cases in ancient egypt

• Addition of paper pulp or cow hair into POP

• Ply Wood

• Reinforced Concrete

Fatigue in Nature• Nature Has Used Very Advanced Fracture Mechanics

Concepts For Designing These Structures

• If It Is Not A Composite, Assume That The Tree Is Made Out Of Isotropic Material

• Any Small Damage Caused By Human Beings Or Animals Would Have Grown To The Critical Size In A Short Time Due To Wind Loads and No Tree Would Have Survived More Than Couple of Months

• Nature Created A Fibrous Structure Where Damages Are Arrested Locally

• Human Beings Took So Long Time To Realize That To Overcome Fatigue,We Have To use Composites Which Nature Realised Many Millions Of Years Back !!!!!

Specific Modulus

Sp

ecif

ic S

tre

ng

th

• Cost of materials and processes are generally high in civil aircraft applications.

• Unlike metals CFRP has very poor post –yield deformation behavior and this makes CFRP a very poor in energy absorption during a failure

• A proper trade off and selection is essential for arriving at the best choice for cost effective and reliable product

• Need to examine positives and negatives as it exists today together with technological steps to overcome problems.

• Boeing 787 and Airbus 350 are two recent cases of maximum composite content .

There are also some limitations

• Composites, depending on the type of resin can be divided into two groups

– Thermoset resin based composites

– Thermoplastic based composites

• Majority of the composites in Aerospace are thermoset resin based composites.

• The production of composite parts with thermosetting matrices involves many steps of which curing is the most critical.

• During curing composite layup is transformed from a soft, multi-layered mixture of resin and fibers to a hard structural component.

Classification

Intro - contd

• During the curing cycle, the resin changes from a liquid to a rubbery and eventually to a solid at the gel point.

• Beyond the gel point, there is little resin movement or flow; typically, the cure is 40 to 60% completed, and the final shape and thickness of the composite are fixed.

• In the curing process, several requirements must be met simultaneously

– consolidation and void elimination

– solvent removal/removal of excess resin

– Complete fiber wet-out

Fabrication Processes for Thermoset Composites

• Hand Layup

• Spray up

• Vacuum Bag Moulding

• Prepreg Manufacture

• Autoclave Moulding

• SMC

• DMC

• Filament Winding

• Pultrusion

• Liquid Composite Moulding

Fabrication of Aerospace Structures

• Since we are dealing with flexible materials, there are many possibilities for making parts

• For aerospace structures, these methods can be divided into two

– Using prepregs

– Using liquid resins

– Using Films

• For prepregs most popular are

– Vacuum bag method

– Autoclave method

Prepreg Manufacture

AUTOCLAVE MOULDING

• The autoclave method has similarities to the vacuum-bag method but has much more sophisticated features.

• It is extensively used in the aerospace industry because of modest tool requirements and the ease of cure control.

• The autoclave method is usually reserved for the processing of prepreg materials with high-performance reinforcements.

Prepreg Layup

• Prepregs are cut to size and laid up either individually or in staples on a tool which is not required to sustain high pressure loads.

• A peel ply and several bleeder plies are placed on top of the prepregs which absorb the excess resin during curing

• Separated by a perforated sheet, several breather plies may be added to provide a suction path in case of soaking of the bleeder plies

• The peel ply is a resin-porous, non-adhering fabric, while the bleeder and breather plies are usually glass fiber weaves.

Prepreg Layup & Curing

• The complete layup is covered with an airtight membrane and sealed against the tool

• The whole assembly is then placed into an autoclave capable of

– drawing a vacuum under the airtight membrane.

– exerting external pressure on the tool surfaces

– delivering the heat necessary for the curing process

• All three parameters need to be controlled accurately.

• Curing process is initiated with the preheating of the autoclave which initiates the resin flow

• The subsequent evacuation of air from under the membrane serves three purposes

– to cause initial compaction of the laminate

Curing

• Curing of the resin is done by elevating the temperature and introducing pressure inside the autoclave

• The temperature and pressure intensities and durations depend on

– type of prepreg material

– dimensions of the part to be produced and

– design requirements

• Cure cycle must be re-established for each new application

• A typical cure cycle for epoxy resins extends to ~ 6 hours.

Post Curing

• Generally Epoxies require a post cure cycle

• Post-cure temperature is slightly greater than the curing temperature

• Effect of Post Curing

– Completes the curing

– improves stiffness

– Increases the Glass Transition temperature

– Decrease in Toughness

Typical Autoclave Cure Cycle

Characteristics of the Autoclave method

• Tool material: Steel, aluminum, reinforced plastics

• Curing temperature: 25 - 300˚ C

• Curing pressure: 1 – 40 bar

Advantages

• Uniform and High quality of parts

• Easy Process Control

Disadvantages

• High initial investment cost

• Need for expert personnel

• Size of Part limited by Autoclave Dimensions

What makes composites attractive ??

50%Weight of Structure

25%Payload

25%Fuel

Composite Light Weight Structure

25%Fuel

45%Weight of Structure

10%30%Payload

20%

Economic Gains

through

Light Weight Construction

Tensile Strength of Al & CFC

45 %

0 %

Tensile Strength & Fatigue Life of Al & CFC

Tensile Strength & Fatigue Life of Al & CFC

Fatigue in the Presence of Damage

Status of Composites

Composite Structures in a Typical Aircraft

Boeing 777

Therefore, in order to reduce production costs and weight, research has been conducted in applying the vacuum assisted resin transfer molding or "VARTM“ technique to the fabrication of the primary composite structures of the MJ aircraft. The aim here is to apply the technique to the fabrication of vertical stabilizer.

Research in the Application of the VaRTM Technique to the Fabrication

of Primary Aircraft Composite Structures- Mitsubishi Heavy

Industries, Ltd.

Usage of composites in Military aircraft

India makes it to Global Composites Scene with LCA- Tejas Program

Composites in A 380

A 350XWB Material Breakdown

DESIGN

&

ANALYSIS

ADVANCED

RESEARCH

STRUCTURAL

HEALTH

MONITORING

STRUCTURAL

REPAIR

MANUFACTURING

NAL’S CORE STRENGTH IN COMPOSITES

NON-

DESTRUCTIVE

EVALUATION

STRUCTURAL

TESTING

PROCESS

DEVELOPMENT

Composites – Present Drivers

• Weight reduction

• Improvements in fatigue resistance

• Corrosion prevention

• Other benefits

Potential fabrication cost advantages for parts with complex

shapes

Performance advantages (e.g., damage tolerance)

Issues to be addressed

• Integration of structural design detail with repeatable

manufacturing processes

Material and process control

• Design details, manufacturing flaws and service damage,

which cause local stress concentration

Strength, fatigue & damage tolerance

Dependency on tests

Scaling issues

• Environmental effects

Temperature

Moisture content

• Maintenance inspection and repair

Data needed for Aerospace Design

How do we Manufacture Composites ???

Courtesy DLR, Germany

Co-curing Vs Co-bonding

Co-curing Co-bonding

All parts in the assembly are cured in

a single cure

Some parts in the assembly are pre-

cured

Single stage fabrication Multi-stage fabrication

Parts count is less Parts count is more

Interfaces between the parts need no

preparation

Interfaces of pre-cured parts need

preparation

Trimming is done only at the edges Trimming need to be done on all pre-

cured parts

Manufacturing cost is less Manufacturing cost is more

Co-curing Vs Co-bonding

Co-curing Co-bonding

Cycle time is less Cycle time is more

Material consumption is less Material consumption is more

Probability of part rejection is higher Probability of part rejection is low

Tools required for manufacturing the

part are quite complex

Tools required for manufacturing the

part are relatively simple

The interface between the sub-

structure and the skin is under tensile

stress because of spring-in

phenomena

This is not the problem with the co-

bonded structure. Compensation for

the spring-in phenomenon can be

done.

Benefits of Integration through Cocuring

• No holes- No stress concentration

• Increased stiffness of structure

• Better aerodynamic surface

• Reduced assembly time

• Weight saving

• No fuel leakage

CSIR-NAL has developed Cocuring technology

within the country for

Light Combat Aircraft (LCA-Tejas) and SARAS aircraft

Challenges

• Curing Distortions-tolerances are very stringent

• Results in complex tool design

• ‘your part is as good as your tool’

Parameters influencing curing distortions• Resin characteristics ---Modulus and co-

efficient of thermal expansion• Fiber properties- Modulus and co-

efficient of thermal expansion• Cure temperature• Fillet radius & Flange thickness

Probable Solutions

• Chamfer the flange such that thickness varies

from root to the end of flange.

• Eliminate the corner resin build-up.

• Predict the deformations and give these

corrections in the outer profile of the skin, so

that after curing it is in the desired shape.

Composite Parts made for LCA-Tejas by NAL

45% by weight in composites

Manufacturing of Cocured Structures

Typical Fin

Co-cured Fin for Tejas (LCA)

Co-cured torsional box

Fin assembly

Torsional box internal details

Root clips

Hinge brackets

Co-cured fin tip rear

Co-cured Fin tip fwd

RWR Antennae

fairing fwd

Sparlets

Co-cured nose box

Root ribRoot fittings

Tip rib

Fairing angle

Co-cured torsional box

Exploded view of Fin assembly

Fairing angle nose

Comparison of CFC vs Metal

Metal CFC

Weight 45.5 kgs 38.5 kgs

No of parts 27 1

No of fasteners

2500 Nil

Assembly time

4 weeks Nil

Highlights

Reduced assembly time

leading to productivity

Smooth aerodynamic

surface

Spars

Nose clips

Mid ribs

Highlights

Integral construction

Fasteners reduction

20% weight saving

Lower cost

INTER SPAR BOX MOULD

INTER SPAR BOX MOULD ASSEMBLY

SPAR LAYUP

SPAR WITH LAY UP TOOLS

SPAR LAY UP TOOLS

EVEN SURFACE ROWINGS TO MAKE FILLED WITH UD

SPAR LAY UP

SECTION-AA

A

A

TOOLS

SPAR LAY UP

UD ROWINGS

SPAR LAY UP

LAY UP TOOLS

EVEN SURFACE

SPAR LAY UP

SPAR LAY UP

TOOLS

SPARS RELEASED 5

FILLED WITH UD ROWINGS TO MAKE

TOOLSSPAR LAY UP

AND DO THE DEBULK

ASSEMBLE THE SPARS43

ASSEMBLE THE SPARS

AND DO THE DEBULK

SPAR LAY UP

SPAR LAY UPTOOLS

TOOLSSPAR LAY UP

SPAR LAY UP

WITH SPAR LAY UP TOOLS

SECTION OF SPAR LAY UP2

1SECTION OF SPAR

SPAR LOCATING TOOLS

SPAR LOCATER ASSEMBLY

Prior to cure

Cured part

Cocured Composite Structures Developed for LCA

Fuselage Top Skin

Centre Fuselage: Trouser Duct-Top

Fin

Rudder Torque Shaft for Rudder

Landing Gear Door

Q Hybrid (metal + composite) airframe

Q CFC flaps, control surfaces, fairings

Q P&WC PT6A-67A turbo-prop engine

l 1200 SHP

Q 2.65 (5 bladed) constant speed propeller

Q Max. cruise speed : 550 km / h

Q Max. cruise altitude : 9 km

Q Max. R/C, ISA, SL : 700 m / min.

Q Endurance : ~ 5h

Q T.O. distance, ISA, SL : 700 m

Q Landing distance, ISA, SL : 850 m

Design to meet FAR-23 requirements

14 seater multi-role LTA - SARAS

Development of a Light Transport Aircraft

Composite Parts Made For SARAS Aircraft

WING

ELEVATOR

INBOARD

FLAPOUTBOARD

FLAPAILERON

HORIZONTAL

STABILIZER

REAR PRESSURE BULK HEAD

FLOOR

BOARD

FUSELAGE TOP SKIN

FIN35% by weight in composites

New Processing

Technology:

VERITy

Nacelle

Comparative chart

Weight

No. of parts

fastenersNo. of

Assembly

32.0 kgs. 24.0 kgs.

75 1

5200 Nil

4 weeks Nil

Metal Composites

Horizontal Tail aft box

Dimensions: 5.5mx1m. The skin is co-cured with stringers, ribs and spars.

Horizontal Tail of SARAS: Cocured Bottom Section

HT Components of SARAS

Metal Composite

Weight 92 Kg 70 Kg (24%)

No. of parts 243 11

No. of Fasteners 10,500 2900

Cocured Inter Spar Box with Bottom Skin With 2 Spars, 11 Ribs, 7 Stringers

Cocured Top Skin with Stringers

HT Tip Cocured with Stringers

Size: 5.5mx 1m

Basic outer CFC Mould Internal Flexible tools

Skin stringer Integration Skin stringer spar Integration

Final bag for curing

Tooling Concepts

Development of Composite Fin for SARAS

FIN ASSEMBLY

DescriptionMetal

Design

Compos

ite

Design

Weight

Reduction

WEIGHT 65 kg 51 kg 22%

NO. OF

PARTS130 01 ———

NO. OF

FASTENE

RS

1100 NIL ———

CO-CURED INTER SPAR BOX (INTERNAL DETAILS)

Spars

Mid Rib

I.S Skin (LH)

I.S Skin (RH)

Nose Box

Co-cured I.S Box

Tip Cap

HT cutout

CO-CURED INTER SPAR BOX

Bottom Fairing

COMPARATIVE CHART FOR CO-CURED IS BOX

Skin lay-up Internal Bag Spar lay-up

Spar Caul PlateRadius Caul PlateMould

UNIFORM PRESSUREAPPLICATION ON THE PART

Master Model For MouldRh & LH Mould Assembly

Skin Bonded With Spars & Mid Ribs

Final Bagging for Curing

Cured Component

Co-cured CFC Pressure Bulkhead

Ring

Dome shaped rear wall

Diameter = 1.8 mDepth = 175 mm

Metal Composite

Weight in Kg 3417

( 50 % saving)

No. of Fasteners 700 0

Nacelle parts for SARAS

Mouldability of Composites Exploited through Innovative Tooling

Technology

Cost!!!

• All the above were fabricated using Prepregs and

Autoclave Moulding Technology

Challenge: How to cut costs???

One solution- Liquid Moulding Technology

Cost breakup for typical aircraft structures

Material cost = 25% and total processing cost is 75%

Challenge – How to reduce these costs?

LCM and its Variants

RTM (Resin transfer moulding)

RIM ( Resin injection moulding)

VARTM ( vacuum assisted resin transfer moulding)

• SCRIMP ( Seeman composite resin infusion moulding process)

• DCVRTM (Double chamber vacuum resin transfer moulding)

• FASTRAC ( Fast remotely activated channels)

RFI ( Resin Film Infusion)

SRIM ( Structural reaction injection moulding)

VERITy ( Vacuum enhanced resin infusion technology) Developed by NAL

New Fabrication Technology:

Vacuum Enhanced Resin Infusion

Technology (VERITy)

Why VERITy?

• Aircraft composite structures predominantly prepreg technology Prepregs are expensive

Outlife restricts size and complexity of component

Infrastructure such as clean room, cold chest etc required

• Can we make high quality composites at a reduced cost? VERITy aims to achieve this (At least 20% Cost Reduction)

No shelf life items

• SARAS Wing development using VERITy

VERITy Process

Vacuum pump

Reinforcement

MouldResin infusion

Resin impregnatesfibers under vacuum

Cured part

Resin

Consolidation Under 1 Bar External Pressure and Vacuum

Vacuum pump

Vacuum pump

Reinforcement

MouldResin infusion

Resin impregnatesfibers under vacuum

Cured part

Resin

Consolidation Under 1 Bar External Pressure and Vacuum

Vacuum pump

Development of Integrated Wing Structures at

NAL using VERITy Process

SARAS Wing: Substructure Details

Flow Sensor Development

Fibre Optic Flow Sensor

Process Sensor – 1 Sensor – 2 Sensor – 3

After Embedment

a a a

Before Infusion

b b b

Resin crossed

Sensor – 1

Resin crossed

Sensor - 2

Resin crossed

Sensor - 3

Instrument

Res

in F

low

S-1 S-2 S-3

c c c

c b b

c c b

180 mm

100 mm

ResinVIEW Software Development• LabVIEW & MATLAB based modular code development for real time resin flow.

• Enables sequential infusion based on NetSense feedback.

• Resin arrival time information important for future infusion strategy and modeling.

• Low cost reusable sensor & modular open system architecture system.

00:00:00 00:54:15 01:31:38

SARAS Outboard Wing: Integrated Wing Concepts

Cocured Coinfused Wing Bottom Skin with Substructure

Cocured

Ribs and

Stringers

Cocured

Rib with

Gussets

Cocured

Spar with

Gussets

@ 300 parts Cocured in one shot

Vacuum Enhanced Resin Infusion Technology (VERITy)

Basic Technology

Research & Feasibility

Technology Demonstration

Sub Components

Technology Development

Generic Features

TRL: 3-6

Product Development

Components

Process Developed

Product Developed

2mm 20mmTRL: 1-4

TRL: 5-7

TRL: 6-8SARAS Composite Wing

2004-2007

2006-2007

2006-2008

2008-Now

Skin Stringer Panel & Spar Splice

Laminate Qualified by C-scan

SARAS Wing Test Box

Comparison Composite Metal

Weight 500 Kg 625 Kg

Fasteners Qty. 15000 45000

Technology Developed

Mitsubishi vision: large one piece test box

Vision 2020

NAL is already There!!

However…

• NAL has done the R&D

• These novel ‘high level cocuring R&D technologies’ need

to be brought to production standards in terms of

Automation in layup

Automation in Assembly

Faster NDE

Meeting Production Rate

Other Advances

Forming

Forming C-sectional parts like CFC Spars and Ribs are being extensively used in

realizing huge aircraft structures like wing, fin etc.

In most of the cases Successive lamination technique is being used to

fabricate such CFC parts

Successive lamination approach is labor intensive, time consuming, limited

scope for automation and increased possibilities for inclusions during layup

due to human errors.

Need to develop an alternate fabrication approach to explore scope for

enhancing the improvements in the above mentioned factors.

92

Flat Prepreg Forming

Prepreg Stack during heating inside the oven

Flat Prepreg stack formed by the application of vacuum

93

Study Parameters

1. Flange angle: The residual stresses that develop in fibre-reinforced

laminates during autoclave processing while the laminate is confined to

the process tool often lead to dimensional changes such as spring-in of

angles and warpage of flat sections.

2. Finishing effect on flange angle

3. Part thickness

4. Time study

94

Time Study

95

Studies at NAL

1.Spring-in

• Spring-in angle is lesser in case of parts fabricated using forming approach

• For lower thickness(2S configuration), difference in spring-in angle is higher

• And spring-in angle differences between the two approaches reduces with

the increased thickness and the difference is least for 5S configuration

2. Thickness

3. Time study

• Fabrication time in case of forming approach is considerably less as

compared to successive lamination technique and is simple to fabricate96

B 787 Fuselage

Laser projection for layup and assembly

Future Trends

Development of 3D Interface Fittings

Fittings: Fatigue Critical, 3D State of Stress

Efforts to increase the 3D properties

Use of 3D Preforms:

How to Characterize?

How to Generate Design Allowables?

Use of Toughened Resin Systems

Typical 3D Preforms

Motivation

•Basic Building Block of Fitting: T Joints

•Weakness: Loads in Out of Plane Direction

A Typical Fitting

Development of a main landing gear attachment fitting using

composite material and resin transfer moulding

Tufting

Tufting

Tufting is an experimental technology to locally reinforce

continuous fibre-reinforced along the Z-direction

• Structural reinforcement

Vertical reinforcement to improve interlaminar shear resistance

Delamination resistance

Reduces cracking or propagation of cracks

• Joining

of Several layers of sub preforms

• Local reinforcement

Tufting of local patches/structures

• Tension free surface

Tufts only remain in position because of frictional force

• Only one side access is required

• Insertion of single needle with single yarn

Tufting technology at the world level

Applications:Airbus A380 Pressure

Bulk headBoeing 787 Landing gear

Braces

Two fold increase in T-Pull out strength

Three fold increase in fracture toughness

Tufting Preliminary studies at CSIR-NAL

Tufting

• Development of co-cured structures is one of the

thrust area for designers

Absence of fibre in third direction

Out of plane loads

Damage/delamintion due to impact

• High damage resistance, pull off load can only propel

the usage of co-cured structures to the next level

• Manual tufting is time taken and doesn't give

repeatability, this requires automation of tufting

technology

Composites - No Limits to Imagination!

Nano Composites