07 Seleccion de Materiales Aeronauticos

Material Selection for Aerospace Applications

Darren Pyfer, P.E.Engineering Specialist Senior

October 16, 2001

10/16/012

Agenda

• Vought Aircraft Industries Corporate Overview

• Material Selection Criteria

• Material Types

• Material Forms

• Examples

Vought Aircraft Industries Corporate Overview

10/16/014

Vought Company Overview

• Largest Single Supplier of Aerostructures to Boeing:

- Producing More of the 747 Structure Than Any Other Commercial Supplier for Boeing

- Producing More of the C- 17 Structure Than Any Other Military Supplier for Boeing

10/16/015

Vought Company Overview (cont.)

• Largest Single Supplier of Aerostructures to Gulfstream Aerospace

– Designed and Build Integrated Wing System for the Gulfstream GV As a Risk- sharing Team Member

10/16/016

Vought Company Overview (cont.)

• Largest Single Supplier of Aerostructures to Northrop on the B- 2 Stealth Bomber Program

– Designed and Built the Intermediate Wing Section of the B- 2 Bomber including the Engine and Landing Gear Bays

10/16/017

Vought Commercial Products

777

GIV HAWKER 800

CF34CF34

CFM56

CF6CF6

GV

767737747

757

10/16/018

Vought Military Products

S-3 F-14 E-2C P-3

V-22

EA-6B

Global Hawk

T-38

C-17 F/A-18E/F E-8C/JSTARS

10/16/019

Vought Product Line Summary

C-17

Empennage Fuselage Doors WingsNacelleComp

ControlSurfaces

GV

737

747

757

767

777

Material Selection Criteria

10/16/0111

Static Strength

• Material Must Support Ultimate Loads Without Failure. Material Must Support Limit Loads Without Permanent Deformation.

– Initial Evaluation for Each Component

– Usually Aluminum Is the Initial Material Selection

– If Aluminum Cannot Support the Applied Load Within the Size Limitation of the Component, Higher Strength Materials Must Be Considered (Titanium or Steel)

– If Aluminum Is Too Heavy to Meet the Performance Requirements, Graphite/Epoxy or Next Generation Materials Should Be Considered

10/16/0112

Stiffness

• Deformation of Material at Limit Loads Must Not Interfere With Safe Operation

– There Are Cases Where Meeting the Static Strength Requirement Results in a Component That Has Unacceptable Deflections

– If That Is the Case, The Component Is Said to Be a ‘Stiffness’ Design

10/16/0113

Fatigue (Crack Initiation)

• The Ability of a Material to Resist Cracking Under Cyclical Loading

– Spectrum Dependant

– Stress Concentration Factors

– Component Is Limited to a Certain Stress Level Based on the Required Life of the Airframe

– Further Processing May Improve Fatigue Properties Such As Shot Peening or Cold Working

10/16/0114

Damage Tolerance (Crack Growth)

• The Ability of a Material to Resist Crack Propagation Under Cyclical Loading

– Slow Crack Growth Design

– Use of Alloys With Increased Fracture Toughness

10/16/0115

Weight

• Low Weight Is Critical to Meeting Aircraft Performance Goals

– Materials Are Tailored for Specific Requirements to Minimize Weight

– Materials With Higher Strength to Weight Ratios Typically Have Higher Acquisition Costs but Lower Life Cycle Costs (i.e. Lower Fuel Consumption)

10/16/0116

Corrosion

• Surface Corrosion

– Galvanic Corrosion of Dissimilar Metals (see Chart)

– Surface Treatments

– Proper Drainage

• Stress Corrosion Cracking

– Certain Alloys Are More Susceptible to Stress Corrosion Cracking (see Chart)

– Especially Severe in the Short Transverse Grain Direction

10/16/0117

Dissimilar Metal Chart

10/16/0118

Stress Corrosion Cracking (SCC) Chart

10/16/0119

Producibility

• Commercial Availability

• Lead Times

• Fabrication Alternatives

– Built Up

– Machined From Plate

– Machined From Forging

– Casting

10/16/0120

Cost

• Raw Material Cost Comparisons

– Aluminum Plate = $2 - $3 / lb.

– Steel Plate = $5 - $10 / lb.

– Titanium Plate = $15 - $25 / lb.

– Fiberglass/Epoxy Prepreg = $15 - $25 / lb.

– Graphite/Epoxy Prepreg = $50 - $100 / lb.

• Detail Fabrication Costs

• Assembly Costs

• Life Cycle Costs

– Cost of Weight (Loss of Payload, Increased Fuel Consumption)

– Cost of Maintenance

10/16/0121

Specialized Requirements

• Temperature

• Lightning and Static Electricity Dissipation

• Erosion and Abrasion

• Marine Environment

• Impact Resistance

• Fire Zones

• Electrical Transparency

10/16/0122

Performance vs. Cost Dilemma

• Highest Performance For The Lowest Cost Is the Goal of Every Airplane Material Selection.

– Mutually Exclusive

– Compromise Is Required

– Define the Cost of Weight to the Aircraft

Material Types

10/16/0124

Aluminum

• Aluminum Accounts for ~80% of the Structural Material of Most Commercial and Military Transport Aircraft

• Inexpensive and Easy to Form and Machine

• Alloys Are Tailored to Specific Needs

• 2000 Series Alloys (Aluminum- copper- magnesium) Are Medium to High Strength With Good Fatigue Resistance but Low Stress Corrosion Cracking Resistance.

– 2024- T3 Is the Yardstick for Fatigue Properties

• 5000 and 6000 Series Alloys Are Low to Medium Strength but Easily Welded

10/16/0125

Aluminum (cont.)

• 7000 Series Alloys (Aluminum- zinc- magnesium-copper) Are High Strength With Improved Stress Corrosion Cracking Resistance but Most Have No Better Fatigue Properties Than 2000 Series

– 7050 and 7075 Alloys Are Widely Used

– 7475 Alloy Provides Higher Fatigue Resistance Similar to 2024- T3

10/16/0126

Aluminum Tempers

10/16/0127

Aluminum Tempers (cont.)

10/16/0128

Aluminum Tempers (cont.)

10/16/0129

Aluminum Comparison Chart

Material Typical Application2024-T3,

T351,T42

High Strength Tension Applications. BestFracture Toughness/Slow Crack Growth Rateand Good Fatigue life. Thick Forms Have LowShort Transverse Properties including StressCorrosion Cracking.

2324-T3 8% Improvement In Strength Over 2024-T3 WithIncreased Fatigue And Toughness Properties.

7075-T6,T651,T7351

High Strength Compression Applications.Higher Strength Than 2024-T3, But LowerFracture Toughness. T7351 has ExcellentStress Corrosion Cracking Resistance andBetter Fracture Toughness Than T6.

7050-T7451 Better Properties Than 7075-T7351 In ThickerSections.

10/16/0130

Titanium

• Better Strength To Weight Ratio Than Aluminum or Steel

• Typically Comprises ~5% By Weight in Commercial Aircraft and Up To ~25% By Weight For High Performance Military Aircraft

• Good Corrosion Resistance

• Good Temperature Resistance

• Good Fatigue And Damage Tolerance Properties In The Annealed Form

• Typical Alloy Is Ti 6Al- 4V Either Annealed or Solution Treated and Aged

• High Cost For Metals

10/16/0131

Steel

• Steel May Be Selected When Tensile Strengths Greater Than Titanium Are Necessary

• Steel Is Usually Limited to a Few Highly Loaded Components Such As Landing Gear

• There Are Many Steel Alloys to Choose From (See Chart); Select the One That Is Tailored for Your Application.

10/16/0132

Steel (cont.)

Mil- Hdbk- 5 List of Aerospace Steel Alloys:

10/16/0133

Composite

• The Embedding of Small Diameter High Strength High Modulus Fibers in a Homogeneous Matrix Material

• Material Is Orthotropic (Much Stronger in the Fiber Oriented Directions)

• Fibers

– Graphite (High Strength, Stiffness)

– Fiberglass (Fair Strength, Low Cost, Secondary Structure)

– Kevlar (Damage Tolerant)

• Matrix

– Epoxy (Primary Matrix Material) to 250° F

– Bismaleimide (High Temp Applications) to 350° F

10/16/0134

Material Properties Comparison

Material Ftu

(ksi)Fty

(ksi)Fcy

(ksi)E(10

6psi)

Density(lb/in

3)

2024-T3 Aluminum 64 47 39 10.5 .1017075-T6 Aluminum 78 71 70 10.3 .1016Al-4V TitaniumAnnealed

134 126 132 16.0 .160

6Al-4V TitaniumSolution Treated andAged

150 140 145 16.0 .160

15-5PH StainlessSteel (H1025)

154 145 152 28.5 .283

Fiberglass Epoxy(Unidirectional)

80 60 5 .065

Graphite Epoxy(Unidirectional)

170 140 22 .056

10/16/0135

Next Generation Materials

• Aluminum Lithium

• GLARE (Fiberglass Reinforced Aluminum)

• TiGr (Graphite Reinforced Titanium)

• Thermoplastics

• Resin Transfer Molding (RTM)

• Stitched Resin Fusion Injected (Stitched RFI)

10/16/0136

Mil-Hnbk-5 Overview

• Document Contains Design Information On The Strength Properties of Metallic Materials and Elements for Aerospace Vehicle Structures. All Information and Data Contained in This Handbook Have Been Coordinated With the Air Force, Army, Navy, Federal Aviation Administration and Industry Prior to Publication and Are Being Maintained As a Joint Effort of the Department of Defense and the Federal Aviation Administration.

10/16/0137

Basis of Properties

• Material Property Selection Is Dependant on the Criticality of the Structural Component

– Critical Single Load Path Structure

– A Basis (99% Probability of Exceeding)

– S Basis (Agency Assured Minimum Value)

– Other Primary Structure With Redundant Load Paths

– B Basis (90% Probability of Exceeding)

– Without a Test, A or S Basis May Be Required

– Secondary Structure

– B Basis (90% Probability of Exceeding)

10/16/0138

Grain Direction

10/16/0139

Material Properties (Mil-Hdbk-5) Example

• Type

Material Forms

10/16/0141

Sheet

• Rolled Flat Metal Thickness Less Than .25”

– Fuselage Skin

– Fuselage Frames

– Rib and Spar Webs

– Control Surfaces

– Pressure Domes

• Good Grain Orientation

• Many Parts and Fasteners

• Fit Problems

– Straighten Operations

– Shims

– Warpage

10/16/0142

Plate

• Rolled Flat Metal Thickness Greater Than .25”

– Wing and Tail Skins

– Monolithic Spars and Ribs

– Fittings

• Unitized Structure; Fewer Fasteners

• Grain Orientation Can Be a Problem

• High Speed Machining Has Lowered Fab Costs

10/16/0143

Extrusion

• Produced By Forcing Metal Through a Forming Die At Elevated Temperature To Achieve The Desired Shape

– Stringers

– Rib and Spar Caps

– Stiffeners

• Grain Is Aligned in The Lengthwise Direction

• Additional Forming and Machining Required

• Used In Conjunction With Sheet Metal Webs

10/16/0144

Forging

• Produced by Impacting or Pressing The Material Into The Desired Shape

– Large Fittings

– Large Frames/Ribs

– Odd Shapes• Control Grain

Orientation

• Residual Stresses Can Cause Warpage

• Tooling Can Be Difficult

10/16/0145

Casting

• Produced By Pouring Molten Metal Into A Die To Achieve The Desired Shape

– Nacelle/Engine Components

– Complex Geometry

• Dramatically Lowers Part and Fastener Counts

• Poor Fatigue And Damage Tolerance Properties

• High Tooling Costs

10/16/0146

Composite

• Produced By Laying Fabric, Laying Tape, Winding, Tow Placement and 3D Weaving or Stitching

– Skins

– Trailing Edge Surfaces

– Interiors and Floors

• Properties Can be Oriented To Load Direction

• Excellent Strength To Weight Ratio

• High Cost Of Material and Processes

• Poor Bearing Strength

Examples

10/16/0148

Upper Wing Cover

• Skin - 7075- T651 Aluminum Plate

• Stringers - 7075- T6511 Aluminum Extrusion

• After Machining; Age Creep Formed To - T7351/- T73511

• Compression Dominated

• Reduces Compressive Yield Strength

• Greatly Increases Stress Corrosion Resistance

10/16/0149

Lower Wing Cover

• Skin - 2024- T351 Aluminum Plate

• Tension Dominated

• Good Ultimate Tensile Strength

• Very Good Fatigue and Damage Tolerance Properties

• Stringers - 7075- T73511 Aluminum Extrusion

• High Ultimate Tensile Strength

• Good Damage Tolerance Properties

10/16/0150

Spars

• 7050- T7451 Aluminum Plate

• High Tensile and Compressive Strength in Thick Sections

• Good Stress Corrosion Resistance

10/16/0151

Fixed Trailing Edge Surface

• Graphite/Epoxy Fabric

• Aramid/Phenolic Honeycomb

• Fiberglass/Epoxy Fabric Corrosion Barrier

• Secondary Structure

• Stiffness Design

10/16/0152

Leading Edge

• 2024- 0 Clad Aluminum

• Heat Treated to - T62 After Stretch Forming to Shape

• Clad For Corrosion Resistance

• Polished For Appearance

• De- icing by Hot Air/Bird Strike Resistance

10/16/0153

Landing Gear Support Beam

• Titanium 6Al- 4V Annealed Forging

• High Strength and Stiffness

• Critical Lug Design

• Height is Limited By Wing Contours

• Annealed Form Is Good For Fatigue And Damage Tolerance

10/16/0154

Wing to Body Attachments

• PH13- 8Mo Cres Steel Bar

• Critical Lug Design

• High Strength Requirement

• Good Corrosion Resistance

10/16/0155

Flap Tracks

• PH13- 8Mo Cres Steel Bar

• Geometry Is Very Limited By Requirement To Be Internal To The Wing

• Results In Very High Stress Levels

• High Stiffness Is Required To Meet Flutter and Flap Geometry Criteria

• Good Corrosion Resistance