Speaker: K S Narayana Rao...

1. COMPOSITES IN AEROSPACE STRUCTURES
K S Narayana Rao
Technical Manger-Aerospace
LORD INDIA PVT.LTD.
Bangalore

2. April 2009
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CONTENTS
• COMPOSITE APPLICATIONS
• DESIGN PROCESS
• MANUFACTURING TECHNIQUES
• QUALIFICATION PROCESS
• HELICOPTER APPLICATION
• OPERATIONAL FEEDBACK
• CONCLUSIONS

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COMPOSITE APPLICATIONS

4. April 2009
COMPOSITE APPLICATIONS
THE DEVELOPMENT OF ADVANCED
COMPOSITE MATERIAL HAS CONSTITUTED A
REVOLUTION IN MATERIALS APPLICATIONS
IN RECENT YEARS.
THE HIGH STIFFNESS AND STRENGTH TO
WEIGHT OF FIBERS, ALONG WITH OTHER
PROPERTIES SUCH AS ENVIRONMENTAL
RESISTANCE, MAKE COMPOSITE MATERIALS
INCREASINGLY POPULAR AS POTENTIAL
CANDIDATE.
COMPOSITE MATERIALS ARE BEING
INCREASINGLY UTILIZED IN MANY FIELDS
INCLUDING BOTH MILITARY AND
AEROSPACE APPLICATIONS, SPORTING
GOODS, AND CHEMICAL INDUSTRIES.

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AEROSPACE COMPOSITE APPLICATIONS
Typical aerospace applications are
window frames, seat pedestals, fittings,
frame gussets, intercostals, pressure
pans and static engine parts.
Structures include wing-to-body fairing assemblies, wing trailing edge
assemblies, tail section components, radomes, dorsal and flap-track fairing
assemblies, cockpit sidewall and ceiling panels and doorliners, leading
edges, access doors and rotor blade components.

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HELICOPTER COMPOSITE APPLICATIONS
Structures include fairing
assemblies, rotor blades and hub,
tail section components, radomes,
doors, cockpit sidewall and ceiling
panels, engine cowlings and
doorliners and access doors.

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COMMERCIAL COMPOSITE APPLICATIONS
Wind Mill Blades Automobile Body
(Chassis and body)

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ACOUSTIC COMPOSITE APPLICATIONS
Acoustical treatment of
aircraft engines
Non-metallic permeable cap material
embedded into honeycomb core to
create an acoustic septum, with the
caps bonded to the core cell wall with
adhesive.
Honeycomb core
Non-metallic
permeable cap

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SPORT COMPOSITE APPLICATIONS
Laminates used predominantly in the
manufacture of Skis and Snowboards
with durable materials are required to
resist high stresses and strains.
Bike with exceptionally precise
carbon fibre shapes without the
need for secondary machining
operations.

10. BOEING 747 VERSUS 787: COMPOSITES

13. 13
COMPOSITE APPLICATION IN HELICOPTER

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DESIGN PROCESS

15. DESIGN PROCESS
DESIGN PROCESS INVOLVES FOLLOWING THE DESIGN GUIDELINES FOR
PREDICTION OF STIFFNESS, STRENGTH AND FATIGUE CHARACTERISTICS.
DESIGN GUIDELINES
• BALANCE AND SYMMETRIC WHEREVER POSSIBLE.
• HIGH STIFFNESS IN THE FIBRE DIRECTION BUT ENSURE DESIGN INCORPORATES
AN UNDERSTANDING OF STRESS TRANSVERSE TO THE FIBRE.
• CHANGE IN FIBRE ANGLE BETWEEN LAMINAE TO A MINIMUM.
• AVOID SHARP CHANGES IN SECTION TO AVOID STRESS CONCENTRATIONS.
• RIBS AND INTEGRAL STIFFENERS HELP STABILISE LARGE FLAT SURFACES
• GOOD ELECTRICAL CONTACT BETWEEN ALL METALLIC AND CARBON/EPOXY
STRUCTURAL COMPONENTS FOR THE DISSIPATION OF STATIC AND LIGHTNING-
INDUCED CURRENTS.
• MINIMISE POTENTIAL GALVANIC CORROSION AND/OR THERMAL EXPANSION
PROBLEMS BY SELECTING COMPATIBLE MATERIALS.
• USE TITANIUM ALLOY FASTENERS OR OTHER MATERIALS THAT ARE COMPATIBLE
WITH CARBON/EPOXY TO PREVENT GALVANIC CORROSION.

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Analysis Test Aspects and Life Evaluation of Helicopter/ Aircraft
Components of Metal and Composites
Helicopter / Aircraft Components
Analysis
Testing Life
Evaluation
Metals Composites
STRENGTH PREDICTION

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HELICOPTER AND AIRCRAFT MATERIAL
A/C and H/C Materials
Metal
Steel Al.
Alloys
Ti.
Alloys
Mg.
Alloys
Nimonics Al. Li.
Alloys
Composite
Fibers Resin
Glass
Kevlar
Carbon
Thermoset
Thermo
plastic

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MATERIAL REQUIREMENT
Stiffness
 E, G, u
Strength
 Static, Fatigue, Impact, Fracture toughness
Manufacturing
 Machinability, Forgings, Castings, Welding
Environmental
 Temp, Humidity, Wear, Fretting, Corrosion
Cost
 Raw material, Manufacturing

19. Material Data
ISOTROPIC: E – YOUNGS MOD.
G – SHEAR MOD.
 – POISSON RATIO
ORTHOPIC: E11, E22 – YOUNG MOD.
 12 – POISSON RATIO
G12 – SHEAR MOD.
UD QUASAI  45°
E11 43400 21730 13500
E22 11500 21730 13500
G12 4336 8324 13400
12 0.27 0.301 0.565
21 0.072 0.301 0.565
GLASS COMPOSITES
UD QUASAI  45°
E11 116500 33050 13000
E22 8000 33050 13000
G12 3500 16750 30000
12 0.32 0.44 0.8
21 0.072 0.44 0.8
CARBON COMPOSITES

20. MATERIAL PROPERTIES FOR ANALYSIS
• Hence from the above stiffness matrix it can be
inferred that for an anisotropic lamina, the following
material constants along with the corresponding
strength values are required for doing analysis:
E11 – Modulus along fiber direction
E22 – Modulus transverse to fiber direction
G12 – Shear Modulus
12 – Poisson’s ratio
t
11 – Tensile strength along fiber direction
t
22 – Tensile strength transverse fiber direction
c
11 – Compressive strength along fiber direction
c
22 – Compressive strength transverse fiber direction
12 – Shear strength (ILSS)

21. COMPARISON TABLE OF MATERIAL PROPERTIES
MATERIAL

N/mm2
E
N/mm2

g/cm3
E/ /
Aluminum 450 70000 2.8 25000 160
Titanium 1000 100000 4.2 23809 238
Steel 1000 210000 7.8 26923 128
Glass 1100 43400 2.06 21067 533
Carbon 1350 130000 2.12 61320 636
Kevlar 1000 83600 1.5 55733 666

22. Loads: Materials: Metallic
Tension Composite: glass/carbon
Compression kevlar
Shear
Bending
Torsion
Combination of the above
Optimisation involves : Geometry for a given loads
Material
Lay-up sequence
OPTIMISATION

23. Examples - Tension
• Tension load: F = 40kN
• Factor of Safety = 1.5
• Length of the member = 450 mm
Material
Tensile
Strength
N/mm2
Area of c/s
mm2
Weight
Kgs
Remarks
Aluminium 400 150.0 0.189
Shape is
immaterial
Steel 900 66.6 0.234
Titanium 900 66.6 0.12
Carbon
composite
1200 50.0 0.075
Glass
composite
1000 60.0 0.114
• Tension load, F = t x A

24. Examples - Compression
• Compression load: F = 40kN
• Factor of Safety = 1.5
• Length of the member = 450 mm
Material
Young’s
Modulus
N/mm2
Moment of
Inertia
mm4
Shape Topology
b d b,t d,t
Aluminium 70000 70345
b = 30.3 mm 34.6 36, 3 42,3.5
A = 918 mm2 940 396 423
W = 1.157 Kg 1.184 0.498 0.532
Steel 210000 23447
23 26.3 27.5, 2.25 32,3
529 543 228 273
1.857 1.907 0.8 0.958
Titanium 100000 49242
27.7 31.64 32, 2.5 36, 3
767 786 295 311
1.381 1.415 0.531 0.56
Carbon
composite
116500
(UD)
42267
26.68 30.46 33, 2.5 38, 3
719 728 305 329
0.48 0.49 0.205 0.222
Glass
composite
40000
(UD)
123105
34.86 39.79 42, 3.5 48, 4
1215 1243 539 552
1.039 1.063 0.46 0.472
• Buckling Load, Pcr = 2EI / 4L2

25. Material
Tensile Strength
N/mm2
Shape Topology
b, h b, h, t b, h, t1, t2
Aluminium 400
b , h = 18, 36 mm 24, 42, 3 24, 42, 4, 2.5
A = 648 mm2 360 277
W = 0.816 Kg 0.454 0.349
Steel 900
14, 27 19, 32, 2.5 18, 32, 3, 2
378 230 166
1.326 0.807 0.583
Titanium 900
14, 27 19, 32, 2.5 18, 32, 3, 2
378 230 166
0.68 0.414 0.298
Carbon composite 1200
12, 24 16, 28, 2 16, 28, 2.5, 2
288 160 126
0.195 0.108 0.085
Glass composite 1000
13, 26 18.5, 31.5, 2.75 19, 32, 2.5, 2
338 245 149
0.289 0.209 0.127
Examples - Bending
• Transverse load: P = 2250 N
• Factor of Safety = 1.5
• Length of the member = 450 mm
• Bending moment, MB = b x Zb
b = Bending Stress, Zb = Section modulus

26. Examples - Torsion
• Torsional load: MT = 400 Nm
• Factor of Safety = 1.5
• Length of the member = 450 mm
Material
Shear Stress
N/mm2
Shape Topology
d d,t
Aluminium 230
d = 23.67 30, 2.5
A = 440 mm2 216
W = 0.792 Kg 0.272
Steel 520
18.0 23, 2
255 132
0.895 0.463
Titanium 520
18.0 23, 2
255 132
0.458 0.238
Carbon composite 350
20.6 26, 2.25
334 168
0.225 0.114
Glass composite 300
21.66 27, 2.5
369 193
0.315 165
• Torsional Load, MT =  x Zt
 = Torsional Stress, Zt = Section modulus

27. Broad steps for Optimum Design
 Define the loads and its criticality at the given cross section.
 Select the optimum geometry for the critical load.
 Select the material and the lay-up sequence for the strength or
stiffness or combination depending the requirement.
 Check the cross section for the combined loads.

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MANUFACTURING TECHNIQUES

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MANUFACTURING TECHNIQUES
VARIOUS COMPOSITE MANUFACTURING TECHNIQUES ARE:
 Vacuum bag moulding
 Autoclave moulding
 Expansion tool moulding
 Resin Transfer molding
 Compression molding
 Injection molding
 Filament winding
 Pultrusion
 Braiding
 Contact Molding

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MANUFACTURING TECHNIQUES
Main Rotor Blade mould with actuators and thermal pads
Platten Press for curingTail Rotor Blade mould
Autoclave

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MANUFACTURING TECHNIQUES
Filament Winding process
Pultrusion Process
Braiding Machine

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QUALIFICATION PROCESS

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QUALIFICATION PROCESS
• THE BUILDING BLOCK APPROACH FOR QUALIFICATION INVOLVES:
• CONSTITUENT TESTING: TO EVALUATES INDIVIDUAL PROPERTIES OF
FIBRES, FIBRE FORMS, MATRIX MATERIALS, AND FIBRE-MATRIX PRE-
FORMS.
• LAMINA TESTING:TO EVALUATES PROPERTIES OF THE FIBRE AND MATRIX
TOGETHER IN THE COMPOSITE MATERIAL FORM. KEY PROPERTIES
INCLUDE LAMINA TENSILE, COMPRESSIVE & SHEAR STRENGTHS AND
MODULI.
• LAMINATE TESTING: THIS CHARACTERISES THE RESPONSE OF THE
COMPOSITE MATERIAL IN A GIVEN LAMINATE DESIGN. KEY PROPERTIES
INCLUDE TENSILE, COMPRESSIVE & SHEAR STRENGTHS AND MODULI,
INTERLAMINAR FRACTURE TOUGHNESS, AND FATIGUE RESISTANCE.
• STRUCTURAL ELEMENT TESTING: THIS EVALUATES THE ABILITY OF THE
MATERIAL TO TOLERATE COMMON LAMINATE DISCONTINUITIES. KEY
PROPERTIES FOR THE AEROSPACE INDUSTRY, INCLUDE OPEN AND FILLED
HOLE TENSILE & COMPRESSIVE STRENGTHS, COMPRESSION AFTER
IMPACT STRENGTH, AND JOINT BEARING AND BEARING BYPASS
STRENGTHS.
• STRUCTURAL SUBCOMPONENT OR FULL-SCALE TESTING: THIS TESTING
EVALUATES THE BEHAVIOR AND FAILURE MODE OF INCREASINGLY MORE
COMPLEX STRUCTURAL ASSEMBLIES WHICH ARE STRUCTURE AND
APPLICATION DEPENDENT.

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COMPOSITE DESIGN IN
HELICOPTER APPLICATION
• AIRFRAME
• ROTOR

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`FUSELAGE & EMPENNAGE
Fuselage Details
Empennage Details

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AIRFRAME ANALYSIS
FE Model of ALH AIRFRAME
FE Model of ALH Empennage
FE Model of ALH for Dynamic
Analysis

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STRUCTURAL TESTS
Static Strength & Fatigue
Tests on Components,
Sub-assemblies & Full
Airframe
MAJOR COMPONENTS
 Metallic & Composite Tail boom
 Horizontal Stabiliser
 Transmission Deck
 Breakaway Fuselage

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“THE ROTORCRAFT MUST BE DESIGNED
TO ENSURE CAPABILITY OF CONTINUED
SAFE FLIGHT AND LANDING (FOR
CATEGORY A) AFTER IMPACT WITH A 2.2
LB (1.0 KG) BIRD AT SPEEDS EQUAL TO
VNE OR VH (WHICHEVER IS LESS) AT
ALTITUDES UP TO 8,000 FEET…”
COMPLIANCE
ONE KG BIRD HIT WAS DEMONSTRATED
ON THE 12 MM THICK WINDSHIELD AT
VELOCITIES OF 260 KMPH
Bird Strike Test on ALH at GTRE, Bangalore
BIRD STRIKE TESTS

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VIBRATION TEST (SHAKE TEST)
• TO DETERMINE NATURAL FREQUENCY,
MODE SHAPE AND DAMPING
• MEASUREMENT OF RESPONSE
CHARACTERISTICS
• TEST SYSTEM CONSISTS OF THE
FOLLOWING:
• 6 CHANNEL EXCITATION
• 256 CHANNEL RESPONSE
MEASUREMENT
• PC BASED DATA ACQUISITION &
ANALYSIS
• IDENTIFICATION OF NATURAL
FREQUENCIES BY MODE INDICATOR
FUNCTION METHOD

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ROTOR DESIGN

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MAIN ROTOR SYSTEM

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MAIN ROTOR BLADE
Aerofoil region Collar SpoonTip
Steel Cap
Carbon / Epoxy Skin
Form Core
Carbon / Epoxy Channel
E-Glass / Epoxy SparLead Mass

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FLAPWISE STIFFNESS
DISTRIBUTION
0
2
4
6
8
10
12
0 1 2 3 4 5 6 7
RADIUS
FLAPWISESTIFFNESS
LEAD-LAG STIFFNESS
DISTRIBUTION
0
5
10
15
20
25
30
35
40
0 1 2 3 4 5 6 7
RADIUS
LEAD-LAGSTIFFNESS
STIFFNESS DISTRIBUTION IN FLAP AND LEAD LAG
DIRECTION OF MAIN ROTOR BLADE OF ALH
MAIN ROTOR DYNAMIC CHARACTERISTICS

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STIFFNESS DISTRIBUTION IN
TORSION ALONG THE RADIUS
MAIN ROTOR DYNAMIC CHARACTERISTICS
TORSIONAL STIFFNESS
DISTRIBUTION
0
1
2
3
4
5
6
7
8
9
0 2 4 6 8
RADIUS
TORSIONAL
STIFFNESS
MAIN ROTOR BLADE
0.0
5.0
10.0
15.0
20.0
25.0
30.0
35.0
40.0
0 10 20 30 40 50 60 70 80 90 100 110 120
Rotor Speed, %
NaturalFreq.(Hz)
I FLAP II FLAP III FLAP
I LAG II LAG I TORSION
1 OMEGA 2 OMEGA 3 OMEGA
4 OMEGA 5 OMEGA 6 OMEGA
MRWT Series13 Series14
Series15 Series17
1
6
5
4
3
2
MAIN ROTOR RESONANCE
DIAGRAM

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MAIN ROTOR BLADE TESTING
 MAIN ROTOR BLADE – STATIC & FATIGUE
 Root & Transition Section
 Constant Section : Tension- Torsion
 Constant Section: Resonance Test
Blade Root section test
Tension-Torsion test
Resonance test

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WHIRL TOWER TESTS
• Dynamic balancing & tracking
of blades, measurement of
rotor performance, loads
• Over speed tests & endurance
tests
MAIN ROTOR SYSTEM TESTING

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COMPONENTS TESTED
MRB
TRB
RADOME
FUSELAGE PANELS
Lightning strike tests on Rotors of ALH at
CABS, Bangalore
High Voltage Test on Radome
Test Setup for Direct Effects Test
on Tail Rotor Blade
High Current Test Wave Form
Applicable current wave forms
LIGHTNING PROTECTION TESTS FOR DIRECT EFFECTS

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GROUND TEST VEHICLE (GTV)
 A HELICOPTER, COMPLETE IN ALL
RESPECTS, IS ANCHORED TO THE
GROUND
 TESTS DONE ON ROTOR DRIVE
SYSTEM AND CONTROL SYSTEM TO
ESTABLISH RELIABILITY
 50 HOURS OF ENDURANCE BEFORE
FIRST FLIGHT
 A TOTAL OF 400 HOURS CARRIED OUT
FOR CERTIFICATION

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FLIGHT TEST

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PT – 1
900 Hours
502 Hours
FIVE PROTOTYPES
FLIGHT TESTED
FOR DEVELOPMENT &
CERTIFICATION PROGRAME
LOGGED MORE THAN
2750 FLIGHT HOURS
PT – 2
PT – A
PT –N
PT –C
FLIGHT TESTS
DEVELOPMENT AND CERTIFICATION

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TO & Landing at highest Helipad
(4750 M ALTITUDE)
SHIP DECK LANDING
SLITHERING UNDER SLUNG LOAD
ALH: TRIALS WITH
INFLATED FLOATATION GEAR
Operational capabilities

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OPERATIONAL FEEDBACK

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OPERATIONAL FEEDBACK
• EVEN THOUGH A SYSTEMATIC APPROACH IS FOLLOWED FOR DESIGN AND
DEVELOPMENT OF COMPONENT / SYSTEM, CERTAIN DEFICIENCIES SURFACE UP
DURING SERVICE OPERATIONS DUE TO VARIOUS FACTORS SUCH AS PRODUCTION
VARIATION, OPERATIONAL HANDLING, ASSEMBLY VARIATIONS OR LAPSE IN
MAINTENANCE.
• THE CHALLENGE IS TO REACH THE ULTIMATE FULFILLMENT OF PRODUCT DESIGN
TO THE SATISFACTION OF THE END USER IN TERMS OF ITS PERFORMANCE, SAFETY,
RELIABILITY, MAINTAINABILITY AND COST-EFFECTIVENESS.
• OPERATIONAL FEEDBACK ON THE ABOVE ASPECTS HELPS IN OPTIMISING THE
PRODUCT’S IN-SERVICE UTILIZATION.
• IN ORDER TO FOCUS ATTENTION ON SUCH CRITICAL ISSUES, A FOOL-PROOF
MECHANISM IS REQUIRED TO COLLECT, COMPILE AND ANALYSE OPERATIONAL
FEEDBACK.
• IN THIS REGARD, A TYPICAL CASE STUDY ON THE MAIN ROTOR BLADE
OPERATIONAL EXPERIENCE AND THE CORRECTIVE ACTION TAKEN FOR THE
PRODUCT IMPROVEMENT IS HIGHLIGHTED.

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MAIN ROTOR BLADE
Delamination Region
 During CT scan low CT density was noticed near collar region on a blade
during 500 hrs service
 Low CT values indicate probable debond in the midplane.

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MAIN ROTOR BLADE – Collar section
(with delamination)

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ANALYSIS
 The influence of the debond in respect of structural
integrity and functionality with regard to blade
dynamics has been analysed.
 The stress analysis has been carried out for the limit
loads for the critical cross section with and without
mid-plane debond.
BASELINE
With debond

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STRESS ANALYSIS OF CROSS SECTION
58
Shear stress on section R = 850 mm
Normal stress (z) on section R = 850 mm
BASELINE
With debond

59. 59
Configurations
Axial
EA
(N)x108
Flap
EIx
(N m2)x104
Lead-lag
EIy
(N m2)x105
Torsional
GJ
(N m2)x104
850R
Without debond 2.170 3.089 2.889 4.199
With full debond 2.066 3.089 2.754 3.89
RESULTS & OBSERVATIONS
Stress
Config 1
(No debond)
N/mm2
Config 2
(Mid plane
debond)
N/mm2
Config 3
(Vertical
Debond)
N/mm2
Config 4
(Mid plane &
Vertical debond)
N/mm2
Allowable
(N/mm2)
z (Normal) 293 302 295
304 Axial =1000
 (Shear) 10 13 12.5
13 ILSS = 70
Stress
Config 1
(No debond)
N/mm2
Config 2
(Midplane debond)
N/mm2
Config 3
(Vertical Debond)
N/mm2
Config 4
(Midplane & Vertical
debond)
N/mm2
Allowables
(N/mm2)
z (Normal) 155 161 167 163 Axial = 260
 (shear) 56 138 60.7 139 Inplane = 300
Stress
Config 1
(No debond)
N/mm2
Config 2
(Midplane debond)
N/mm2
Config 3
(Vertical Debond)
N/mm2
Config 4
(Midplane & Vertical
debond)
N/mm2
Allowables
(N/mm2)
z (Normal) 101 105
102
106 Axial = 170
 (Shear) 87 98
111
99 Inplane = 320
Stress levels in spar for limit loads Stress levels in the collar cap for limit load
Stress levels in the skin for limit load

60. Conclusion
From the stress analysis of MR blade collar section following observation are made:
 Influence on the stress level due to mid plane & vertical debond on section R = 850
mm is:
No significant change in normal and shear stress in the spar and skin.
No significant change in normal stress in collar cap, however there in an increase
in shear stress.
 The maximum stresses in spar, skin and collar cap obtained on fully debonded
configuration for limit load are within the material allowable indicating the structural
integrity of collar region of the blade is not jeopardised.
 Change in the stiffness of the cross section due to debond between station 710 R to
850 R is small.
 Based on Analysis and Testing, Blades with Horizontal and Vertical/slant/ lines
observed in CT scan can be allowed to operate further without affecting flight safety
until perceivable delamination in collar /skin area is noticed.

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CONCLUSIONS
• A SYSTEMATIC BUILDING BLOCK APPROACH HAS BEEN EVOLVED OVER THE YEARS FOR
DESIGN AND QUALIFICATION OF COMPOSITE PARTS CONSIDERING THE MATERIAL,
MANUFACTURING PROCESS AND QUALITY CONTROL.
• THE COMPOSITE COMPONENTS HAVE FOLLOWED THE BUILDING BLOCK APPROACH FOR
DESIGN AND QUALIFICATION.
• EVEN THOUGH A SYSTEMATIC APPROACH IS FOLLOWED FOR DESIGN AND DEVELOPMENT
OF COMPONENT / SYSTEM, CERTAIN DEFICIENCIES SURFACE UP DURING SERVICE
OPERATIONS DUE TO VARIOUS FACTORS SUCH AS PRODUCTION VARIATION, OPERATIONAL
HANDLING, ASSEMBLY VARIATIONS OR LAPSE IN MAINTENANCE.
• THE COMPOSITE DESIGN CHALLENGE AND WAY AHEAD IS TO HAVE AN INTEGRATED
APPROACH IN THE DISCIPLINE OF MATERIAL SELECTION, ANALYSIS, MANUFACTURING
PROCESS, QUALITY CONTROL AND MAINTENANCE. EVERY DISCIPLINE NEEDS TO
COMPLIMENT EACH OTHER TO OVERCOME THE LIMITATION AND REACH TOWARDS THE
EFFICIENT UTILIZATION OF COMPOSITES TO KEEP THE OVERALL LIFE CYCLE COST TO A
MINIMUM.

62. Thank you