Jet Engine Materials A quick overview of the

Jet Engine Materials A quick overview of the materials requirements, the materials being used, and the materials being developed

Motivation for Materials Development u Higher Operating Temperatures u Higher Rotational Speeds u Lower Weight Engine Components u Longer Operating Lifetime u Decreased Failure Occurrence u This all adds up to: u. Better Performance u. Lower Life Cycle Costs

Materials Requirements u thousands of operating hours at temperatures up to 1, 100°C (2000 °F) u high thermal stresses caused by rapid temperature changes and large temperature gradients u high mechanical stresses due to high rotational speeds and large aerodynamic forces u low- and high-frequency vibrational loading u oxidation u corrosion u time- , temperature- and stress-dependent effects such as creep, stress rupture, and high- and low-cycle fatigue.

Regions of the Engine u Cold Sections u Inlet/Fan u Compressor u Casing u Hot Sections u Combustor u Turbine/Outlet

Cold Section Materials Requirements u High Strength (static, fatigue) u High Stiffness u Low Weight u Materials: u Titanium Alloys u Aluminum Alloys u Polymer Composites u Titanium intermetallics and composites

Applications of Polymer Composites

Fiber Reinforced Polymer Composite Properties Graphite/Kevlar u Very high strength-weight ratios u Very high stiffness-weight ratio (graphite) u Versatility of design and manufacture u Specific gravity: ~1. 6 (compared to 4. 5 for titanium & 2. 8 for aluminum) u Can only be used at low temperatures < 300 °C (600 °F)

Titanium alloys used for critical cold section components u Fan disks/blade u Compressor disks/blades u Typical Alloy: Ti-6 Al-4 V

Titanium Properties u High strength & stiffness to weight ratios > 150 ksi, E = 18 Msi u Specific gravity of 4. 5 ( 58 % that of steel) u Titanium alloys can be used up to temperatures of ~ 590 °C (1100 °F) u Good oxidation/corrosion resistance (also used in medical implants) u High strength alloys hard to work therefore many engine components are cast

Metallurgy of disks critical to achieve desired properties and to eliminate defects u Accident occurred JUL-19 -89 at SIOUX CITY, IA Aircraft: MCDONNELL DOUGLAS DC-10 -10, Injuries: 111 Fatal, 47 Serious, 125 Minor, 13 Uninjured. u A FATIGUE CRACK ORIGINATING FROM A PREVIOUSLY UNDECTECTED METALLURGICAL DEFECT LOCATED IN A CRITICAL AREA OF THE STAGE 1 FAN DISK THAT WAS MANUFACTURED BY GENERAL ELECTRIC AIRCRAFT ENGINES. THE SUBSEQUENT CATASTROPHIC DISINTEGRATION OF THE DISK RESULTED IN THE LIBERATION OF DEBRIS IN A PATTERN OF DISTRIBUTION AND WITH ENERGY LEVELS THAT EXCEEDED THE LEVEL OF PROTECTION PROVIDED BY DESIGN FEATURES OF THE HYDRAULIC SYSTEMS THAT OPERATED THE DC-10'S FLIGHT CONTROLS.

Aluminum alloys can reduce weight over titanium u Conventional alloys have lower strength/weight ratios than Ti but more advanced alloys approach that of Ti. u Specific gravity: 2. 8 ( 62 % that of Ti) u Lower cost than Ti u Max temp for advanced alloys: ~ 350 °C (600 °F) u Lower weight & rotating part inertia

Titanium Aluminide Ti 3 Al u An intermetallic alloy of Ti and Al u Extends the temperature range of Ti from 1100 °F to 1200 -1300 °F u Suffers from embrittlement due to exposure to atmosphere at high temperature - needs to be coated.

Titanium Composites (MMC) u Titanium matrix with Si. C fibers u Decreases weight while increases strength and creep strength TYPICAL Ti/Si. C COMPOSITE 100 X

Hot Section Materials Requirements u High Strength (static, fatigue, creep-rupture) u High temperature resistance 850 °C - 1100 °C (1600 °F - 2000 °F) u Corrosion/oxidation resistance u Low Weight

High Temperatures - 1100 °C (2000 °F) u Creep becomes at factor for conventional metals when the operating temperature reaches approximately 0. 4 Tm (absolute melting temp. ) u Conventional engineering metals at 1100 °C: u. Steel u. Aluminum u. Titanium u Conclusion: ~0. 9 Tm ~1. 4 Tm ~0. 7 Tm We need something other than conventional materials!

High Temperatures - 1100 °C (2000 °F) What Materials Can Be Used? u. Unconventional or superalloys u. Ceramics metal alloys -

Superalloys u Nickel (or Cobalt) based materials u Can be used in load bearing applications up to 0. 8 Tm - this fraction is higher than for any other class of engineering alloys! u High strength /stiffness u Specific gravity ~8. 8 (relatively heavy) u Over 50% weight of current engines

Typical Compositions of Superalloys CHEMICAL COMPOSITION, WEIGHT PERCENT Chromium yields corrosion resistance

Microstructure of a Superalloy u Superalloys are dispersion hardened u Ni 3 Al and Ni 3 Ti in a Ni matrix u Particles resist dislocation motion and resist growth at high temperatures

Creep - Rupture u Strain increases over time under a static load - usually only at elevated temperatures (atoms more mobile at higher temperatures) u The higher energy states of the atoms at grain boundaries causes grain boundaries particularly ones transverse to load axis - to creep at a rate faster than within grains u Can increase creep-rupture strength by eliminating transverse grain boundaries

Controlled grain structure in turbine blades: Equi-axed Directionally Single Crystal solidified (DS) (SX)

Performance of superalloy parts enhanced with thermal barrier coatings u Thin coating - plasma sprayed u MCr. ALY coating materials u Increased corrosion/oxidation resistance u Can reduce superalloy surface temperature by up to 40 °C (~100 °F)

Non-metallics - Ceramics • • • SUPERALLOY Cobalt Nickel Chromium Tungsten Tantalum • Silicon • Nitrogen • Carbon CERAMIC

Ceramics - Advantages u Higher Temperatures u Lower Cost u Availability of Raw Materials u Lighter Weight u Materials: u Al 2 O 3, Si 3 N 4, Si. C, Mg. O

Ceramics - Challenges DUCTILITY IMPACT Superalloys Ceramics TOUGHNESS CRITICAL FLAW SIZE

Ceramic Composites u Ceramic Fiber Reinforced Ceramic Matrix u Improve toughness u Improve defect tolerance u Fiber pre-form impregnated with powder and then hotpressed to fuse matrix

Carbon-Carbon composite u Carbon fibers in a carbon matrix u Has the potential for the highest temperature capability > 2000 °C (~4000 °F) u Must be protected from oxidation (e. g. Si. C) u Currently used for nose-cone for space shuttle which has reentry temperatures of 1650 °C (3000 °F)

TURBINE ROTOR INLET TEMP, F Trends in turbine materials

Materials for F 109 engine F 109 FAN MODULE MATERIALS

F 109 HP COMPRESSOR MATERIALS 201 -T 6 Aluminum INCO 625 (side plates) INCO 718 (vanes) 17 -4 PH Ti 6 -4 INCO 718 Ti 6 -2 -4 -2 HAST X

F 109 COMBUSTOR/MIDFRAME MATERIALS HS 188 + TBC INCO 718 HS 188 HAST S HAST X 300 SS INCO 718 INCO 600

F 109 HP TURBINE MATERIALS INCO 738 HAST X INCO 718 MAR-M 247 HAST S DS MAR-M 247 DS WASP B MAR-M 247 INCO 738 DS

F 109 LP TURBINE MATERIALS HAST X BACK WITH HAST X 0. 032 CELL. HONEYCOMB INCONEL 625 EQUIAXED MAR-M 247 COATED WITH RT-21 INCONEL 625 HAST X BACK WITH HAST X 0. 032 CELL. HONEYCOMB HASTELLOY X WASPALOY EQUIAXED MAR-M 247 COATED WITH RT-21 HAST X BACK WITH HAST X 0. 032 CELL. HONEYCOMB WASPALOY HAST X BACK WITH HAST X 0. 032 CELL. HONEYCOMB WASPALOY