Post on 05-May-2020
transcript
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