Ceramics in Aerospace: Thermal Protection & Engine Applications

Aerospace ceramics do not behave like pottery clay. They are engineered materials that survive conditions no metal alloy can tolerate, including surface temperatures above 3,000°F (1,650°C) during atmospheric reentry and combustion chamber pressures exceeding 2,000 psi.

This guide covers the full spectrum of ceramics in aerospace: thermal protection systems (TPS), ceramic matrix composites (CMC), thermal barrier coatings (TBC), silicon nitride and silicon carbide engine components, ultra-high-temperature ceramics (UHTC), and the emerging role of additive manufacturing in next-generation aerospace ceramic production.

What Are Aerospace Ceramics and How Do They Differ from Conventional Ceramics?

Aerospace ceramics are a subcategory of advanced technical ceramics engineered for extreme thermal, mechanical, and chemical environments. They differ from traditional pottery ceramics in every measurable property: density, fracture toughness, thermal conductivity, oxidation resistance, and operating temperature ceiling.

Traditional stoneware fired to cone 10 (2,381°F / 1,305°C) achieves a modulus of rupture around 3,000 to 5,000 psi. Silicon carbide (SiC) components used in turbine engines operate above 2,500°F (1,370°C) and achieve flexural strength above 400 MPa (58,000 psi), a difference of more than an order of magnitude.

Aerospace ceramics fall into four broad families. The first is monolithic advanced ceramics: single-phase materials such as alumina (Al2O3), silicon nitride (Si3N4), and silicon carbide (SiC). The second is ceramic matrix composites (CMC): ceramic fibers embedded in a ceramic matrix, primarily SiC/SiC systems. The third is thermal barrier coatings (TBC): thin ceramic layers deposited on metal substrates to reduce heat transfer. The fourth is ultra-high-temperature ceramics (UHTC): hafnium and zirconium borides and carbides rated above 3,000°F (1,650°C).

According to the Journal of the American Ceramic Society, the defining attribute that separates aerospace-grade ceramics from industrial ceramics is the combination of low density (2.5 to 3.2 g/cm³ for SiC versus 7.9 g/cm³ for steel) and retained mechanical strength above 1,800°F (982°C), the temperature range where all common metal alloys lose structural integrity.

For readers exploring how advanced ceramics fit into the broader ceramic materials landscape, our overview of ceramic categories from traditional to advanced technical materials provides context for how aerospace grades relate to industrial and structural ceramic families.

The core distinction is this: conventional ceramics are shaped and fired to achieve aesthetic or functional utility at room temperature. Aerospace ceramics are engineered at the atomic level to retain specific mechanical, thermal, and chemical properties under continuous extreme loading.

How Do Thermal Protection Systems Use Ceramics to Survive Reentry?

Thermal protection systems (TPS) protect spacecraft structures from aerodynamic heating during atmospheric reentry. Surface temperatures on the Space Shuttle orbiter’s leading edges reached 2,300°F (1,260°C) on the wing undersurface and exceeded 3,000°F (1,650°C) on the nose cap and leading edge panels, temperatures that would melt aluminum in seconds.

NASA’s Space Shuttle TPS used four distinct ceramic and ceramic-composite material zones, each matched to the peak heat flux at that location on the vehicle. The nose cap and wing leading edges used reinforced carbon-carbon (RCC) composite, a carbon fiber reinforced carbon matrix material with a silicon carbide oxidation-protection coating applied by chemical vapor deposition (CVD).

Reinforced Carbon-Carbon (RCC): The Hottest Zone Material

Reinforced carbon-carbon composites survive above 3,000°F (1,650°C) without structural degradation. Carbon’s sublimation point is approximately 6,500°F (3,593°C), making it the only structural material that strengthens slightly with increasing temperature up to around 4,500°F (2,480°C) due to improved fiber-matrix crystallinity.

The mechanism works because carbon-carbon composites transfer thermal energy through the matrix via phonon conduction rather than electron conduction, which gives them a lower thermal conductivity than metals (6 to 10 W/m·K versus 50 W/m·K for steel). Lower thermal conductivity means less heat reaches the underlying aluminum structure during the 15 to 20-minute reentry window.

Key Specifications for RCC used on Space Shuttle leading edges:

• Operating temperature ceiling: above 3,000°F (1,650°C)
• Density: 1.6 to 1.8 g/cm³
• Thermal conductivity: 6 to 10 W/m·K (parallel to fiber axis)
• Flexural strength: 100 to 140 MPa at 2,200°F (1,200°C)
• Oxidation protection: CVD silicon carbide coating, 0.5 to 2 mm thickness

The failure mode for RCC is oxidation of the carbon matrix at temperatures above 900°F (480°C) when the SiC coating is breached. A cracked or spalled SiC coating exposes bare carbon fiber to oxygen, causing catastrophic oxidative thinning at the rate of 0.1 to 0.3 mm per reentry cycle without repair. This was the mechanism that caused the Space Shuttle Columbia disaster in 2003: a foam impact breached the leading edge RCC panel, and the exposed carbon-carbon oxidized during reentry.

High-Temperature Reusable Surface Insulation (HRSI) Tiles: The Black Tiles

The black tiles covering the Space Shuttle’s underside were made of LI-900, a rigid silica fiber ceramic insulation with 94% open porosity. LI-900 stands for Lockheed Insulation with a density of 9 lbs/ft³ (144 kg/m³), making it one of the least dense rigid ceramic materials ever produced for structural application.

LI-900 works because the 94% void space filled with air provides extreme thermal resistance. Air has a thermal conductivity of 0.025 W/m·K. The silica fiber skeleton that holds the tile together occupies only 6% of the volume, so the effective thermal conductivity of the assembled tile is only 0.02 to 0.06 W/m·K, far below any solid ceramic.

Key Specifications for LI-900 HRSI tiles:

• Density: 9 lbs/ft³ (144 kg/m³)
• Maximum use temperature: 2,300°F (1,260°C)
• Thermal conductivity: 0.02 to 0.06 W/m·K (temperature-dependent)
• Compressive strength: 200 to 600 psi (1.4 to 4.1 MPa)
• Surface coating: reaction-cured glass (RCG), a borosilicate glass-ceramic with SiB4 emittance agent

The condition under which LI-900 fails is water absorption. The open pore structure that gives the tile its insulating value also allows moisture ingress. A saturated LI-900 tile weighs 40% more than a dry tile, shifts the orbiter’s center of mass, and conducts heat far faster. NASA developed a densified variant called LI-2200 (22 lbs/ft³ / 352 kg/m³) for areas requiring higher compressive strength and reduced moisture absorption.

These silica fiber TPS tiles are a category of advanced oxide ceramics used in extreme thermal environments, distinct from the dense structural ceramics used in engine hot sections.

FRCI-12 and Advanced Flexible Reusable Surface Insulation (AFRSI)

FRCI-12 (Fibrous Refractory Composite Insulation) replaced many LI-900 tiles on later Shuttle missions by blending alumina-borosilicate fibers (Nextel 312) with silica fibers in a 20/80 ratio. This blend increased tensile strength by 50% and reduced installation labor by using larger, fewer panels.

AFRSI blankets covered the upper fuselage of the orbiter. They consisted of a high-purity silica fiber batting sewn between a woven silica outer fabric and a beta-cloth inner fabric. AFRSI blankets handled surface temperatures up to 1,200°F (650°C) and replaced approximately 3,000 individual LI-900 tiles per orbiter, reducing maintenance time per flight significantly.

The current generation of reusable launch vehicles, including SpaceX Starship, uses TUFI (Toughened Unipiece Fibrous Insulation) tiles with a densified surface layer that bonds directly to the tile body, eliminating the separate borosilicate coating that cracked and spalled on original Shuttle tiles under high thermal cycling rates.

The bottom line on TPS ceramics: the specific material selected for each zone of a reentry vehicle depends on the peak heat flux at that location, not on a single universal material, and every TPS ceramic is engineered around a specific failure mode that must be prevented over a defined number of reentry cycles.

What Role Do Ceramic Matrix Composites Play in Jet Engine Hot Sections?

Ceramic matrix composites (CMC) are the most commercially significant aerospace ceramic technology in production today. GE Aviation’s LEAP engine, certified for the Boeing 737 MAX and Airbus A320neo, uses CMC components in the high-pressure turbine (HPT) shroud, replacing nickel superalloys that required intensive internal cooling.

CMC is a composite where silicon carbide fibers (typically Hi-Nicalon or Sylramic-iBN grade) are embedded in a silicon carbide matrix. The fiber reinforcement solves the primary weakness of monolithic SiC: low fracture toughness (2 to 3 MPa·m^0.5 for monolithic SiC versus 20 to 25 MPa·m^0.5 for SiC/SiC CMC). When a crack initiates in the matrix, it deflects along the fiber-matrix interface rather than propagating catastrophically.

How CMC Components Are Made: Chemical Vapor Infiltration and Melt Infiltration

CMC hot section components are produced by one of two dominant processes. Chemical vapor infiltration (CVI) deposits SiC into a preform of woven SiC fiber cloth by flowing methyltrichlorosilane (MTS) gas through the fiber architecture at 1,800°F to 2,000°F (982°C to 1,093°C) for 100 to 400 hours. Melt infiltration (MI) infiltrates molten silicon into a carbon-containing preform, reacting in situ to form SiC at the infiltration front.

CVI produces CMC with a cleaner fiber-matrix interface and higher oxidation resistance but leaves 10 to 15% residual porosity. MI produces near-zero porosity CMC with higher thermal conductivity (18 to 25 W/m·K versus 5 to 10 W/m·K for CVI) but requires a residual free-silicon phase that limits maximum operating temperature to approximately 2,400°F (1,315°C).

Key Specifications for SiC/SiC CMC used in GE LEAP HPT shroud:

• Density: 2.7 to 3.0 g/cm³ (versus 8.2 g/cm³ for René N5 nickel superalloy)
• Maximum continuous operating temperature: 2,400°F (1,315°C)
• Tensile strength: 250 to 350 MPa at room temperature
• Interlaminar shear strength: 40 to 60 MPa
• Thermal conductivity: 18 to 25 W/m·K (MI process)
• Weight saving vs nickel alloy equivalent: approximately 20 to 25% per component

The mechanism behind CMC’s temperature advantage is straightforward. Nickel superalloy turbine blades require film cooling: compressor bypass air flows through internal cooling channels, which limits how hot the combustion gases can run. CMC shrouds tolerate 300°F (167°C) higher gas temperatures without cooling air, allowing the engine to operate at higher turbine inlet temperature (TIT), which directly increases thermodynamic efficiency.

According to GE Aviation’s published LEAP engine data, CMC HPT shrouds reduced cooling air consumption by approximately one-third compared to the previous nickel alloy shroud design. This cooling air reduction improved specific fuel consumption (SFC) by approximately 1%, a significant gain across a fleet of thousands of aircraft over decades of service.

The failure mode for SiC/SiC CMC is water vapor oxidation. At operating temperatures above 1,800°F (982°C) in the presence of water vapor (which is abundant in combustion gases), SiC oxidizes to form a silica (SiO2) scale that then reacts with water vapor to form volatile silicon hydroxide (Si(OH)4), which evaporates. This process, called recession, gradually removes material from the CMC surface at 0.001 to 0.01 mm per 1,000 operating hours without a protective coating.

The fix is an environmental barrier coating (EBC): a multilayer ceramic coating deposited by air plasma spray or electron beam physical vapor deposition (EB-PVD) that seals the CMC surface from water vapor. Current EBC systems use a bond coat of silicon metal, a mullite intermediate layer, and a barium-strontium-aluminosilicate (BSAS) or ytterbium disilicate (Yb2Si2O7) topcoat.

CMC Turbine Blades: The Next Frontier

GE Aerospace has demonstrated CMC turbine blades in the GE9X engine (used on the Boeing 777X) and has published a roadmap for CMC blade production at scale. CMC blades are more difficult to manufacture than CMC shrouds because they require complex three-dimensional fiber architectures, internal cooling channel formation, and EBC application to curved aerodynamic surfaces.

The weight advantage of CMC blades over nickel superalloy blades is approximately 1/3 by mass. In a two-spool high-bypass turbofan, each pound removed from the HPT disk and blade assembly reduces the required disk retention hardware by a cascading factor, so a 40-blade HPT stage saving 15 lbs total in blade mass may reduce the total rotating assembly weight by 50 to 70 lbs.

Rolls-Royce and Safran are pursuing parallel CMC turbine blade development programs. According to Safran’s published technical roadmap, CMC blades are targeted for certification in narrow-body engines in the next commercial engine generation, aiming for a turbine inlet temperature increase of 100°F to 200°F (55°C to 110°C) above current nickel superalloy limits.

CMC technology represents the single largest near-term performance gain available to commercial aviation, and the materials science underpinning it is more closely related to industrial SiC ceramics than to any metallurgical tradition.

How Do Thermal Barrier Coatings Protect Metal Engine Components?

Thermal barrier coatings (TBC) are thin ceramic layers deposited on nickel superalloy turbine blades and vanes to reduce the temperature the metal experiences during combustion. A 7 to 8 wt% yttria-stabilized zirconia (YSZ) TBC at 100 to 300 micrometers (0.1 to 0.3 mm) thickness reduces blade metal temperature by 100°F to 300°F (55°C to 167°C), which doubles or triples blade service life due to the exponential relationship between nickel superalloy creep rate and temperature.

YSZ works as a thermal barrier because zirconia (ZrO2) has one of the lowest thermal conductivities among oxide ceramics: 2.0 to 2.5 W/m·K in the dense form, reduced to 1.5 to 2.0 W/m·K in the columnar microstructure produced by EB-PVD. The mechanism is phonon scattering: yttrium oxide (Y2O3) additions stabilize the tetragonal phase of zirconia and introduce oxygen vacancies that scatter phonons, reducing thermal conductivity by 30 to 50% compared to unstabilized ZrO2.

Yttria-Stabilized Zirconia: The Current Standard

7 to 8 wt% YSZ has been the aerospace TBC standard since the 1970s. It is deposited by two methods. Air plasma spray (APS) produces a lamellar (splat-on-splat) microstructure with horizontal cracks that run parallel to the coating surface. EB-PVD produces a columnar microstructure with vertical gaps between columns that grow perpendicular to the substrate surface.

The columnar EB-PVD microstructure is used on turbine blades because vertical column gaps allow the TBC to accommodate thermal expansion mismatch between the zirconia (CTE of 10 to 11 × 10⁻⁶/°C) and the nickel superalloy substrate (CTE of 13 to 16 × 10⁻⁶/°C). If the TBC cannot strain during thermal cycling, it spalls. APS coatings on blades spall after 500 to 1,000 cycles. EB-PVD coatings on blades survive 3,000 to 5,000 cycles.

Key Specifications for 7 wt% YSZ TBC (EB-PVD):

• Thermal conductivity: 1.5 to 2.0 W/m·K (columnar microstructure)
• Maximum use temperature: 2,010°F to 2,190°F (1,100°C to 1,200°C)
• Deposition thickness: 100 to 300 micrometers on blades
• Coefficient of thermal expansion (CTE): 10 to 11 × 10⁻⁶/°C
• Service life on HPT blades: 3,000 to 5,000 thermal cycles
• Bond coat: MCrAlY (nickel-cobalt-chromium-aluminum-yttrium), 75 to 150 micrometers

The failure mode for YSZ TBC is thermally grown oxide (TGO) thickening. The MCrAlY bond coat oxidizes at temperature to form a protective alumina (Al2O3) scale (the TGO) at the metal-ceramic interface. This TGO is beneficial at 0.5 to 2 micrometers thickness because it bonds the YSZ to the substrate. Above 5 to 7 micrometers, the TGO generates stress that drives delamination cracking along the TBC-TGO interface. Spallation follows.

The fix is to monitor TGO thickness using nondestructive methods such as photoluminescence spectroscopy (PLS), which measures chromium ion luminescence shifts in the alumina TGO to estimate thickness without removing the coating. GE Aviation uses PLS inspection on HPT first-stage blades at every engine shop visit.

Next-Generation TBC Materials: Beyond YSZ

YSZ degrades above 2,190°F (1,200°C) by undergoing a phase transformation from the metastable tetragonal phase to a mixture of cubic and monoclinic phases. This transformation produces a 4% volume change that cracks and spalls the coating. As turbine inlet temperatures push above 2,400°F (1,315°C) in next-generation engines, YSZ cannot be used as the sole TBC material.

Research programs at NASA Glenn Research Center, DLR (German Aerospace Center), and multiple university laboratories are evaluating replacement TBC materials. The leading candidates include gadolinium zirconate (Gd2Zr2O7), which is phase-stable to 2,550°F (1,400°C) but has lower fracture toughness than YSZ (1.2 versus 2.0 MPa·m^0.5). Another candidate is lanthanum zirconate (La2Zr2O7), which offers lower thermal conductivity (1.5 W/m·K) and phase stability to 2,730°F (1,500°C). A third candidate is rare-earth aluminates such as gadolinium aluminate garnet (Gd3Al5O12), which offers CTE values closer to nickel superalloys.

Multilayer TBC architectures are emerging as the practical near-term solution: a gadolinium zirconate top coat for high-temperature phase stability combined with a conventional YSZ inner layer for fracture toughness and bond coat compatibility.

For engineers comparing TBC ceramics to structural ceramic systems, understanding how silicon nitride performs in high-temperature oxidizing environments provides essential context, and our detailed technical comparison of silicon nitride properties versus alumina and silicon carbide covers the oxidation resistance mechanisms directly relevant to this comparison.

The practical summary for TBC selection: 7 wt% YSZ applied by EB-PVD remains the production standard for turbine blades below 2,190°F (1,200°C) metal surface temperature, and any new TBC system must match its fracture toughness while exceeding its phase stability to earn a place on a certified engine part.

What Are Ultra-High-Temperature Ceramics (UHTC) and Where Are They Used?

Ultra-high-temperature ceramics (UHTC) are a class of refractory materials with melting points above 5,400°F (3,000°C) designed for hypersonic vehicle leading edges, scramjet combustor liners, and planetary probe nose caps where no other material survives the combined oxidizing and ablative environment. The primary UHTC materials in aerospace development are hafnium diboride (HfB2), zirconium diboride (ZrB2), hafnium carbide (HfC), and tantalum carbide (TaC).

HfB2 has a melting point of 5,990°F (3,310°C) and forms a protective hafnium oxide (HfO2) scale above 3,270°F (1,800°C) that resists further oxidation. ZrB2 has a melting point of 5,630°F (3,110°C) and forms a ZrO2-B2O3 dual-oxide scale at temperatures above 2,730°F (1,500°C). The B2O3 component of the ZrB2 scale is liquid above 842°F (450°C), which allows it to flow and heal surface cracks, a self-sealing mechanism absent in most structural ceramics.

UHTC Composites: Addressing Brittleness

Monolithic HfB2 and ZrB2 have fracture toughness values of 3 to 5 MPa·m^0.5, too low for components subjected to the thermal shock loading of hypersonic flight entry. Research programs at NASA Ames Research Center and the Air Force Research Laboratory (AFRL) have developed UHTC composites that add silicon carbide particles or fibers (typically 20 vol% SiC) to the boride matrix.

The SiC addition improves fracture toughness to 5 to 8 MPa·m^0.5 by crack deflection, increases high-temperature strength by pinning grain boundary sliding, and improves oxidation resistance by forming a continuous SiO2 glassy layer that seals the surface above 2,700°F (1,480°C). ZrB2-SiC composites with 20 vol% SiC retain flexural strength of 350 to 500 MPa at 3,000°F (1,650°C), making them the most capable structural material currently available for hypersonic nose cap applications.

Key Specifications for ZrB2-20vol%SiC UHTC composite:

• Melting point of ZrB2 phase: 5,630°F (3,110°C)
• Maximum demonstrated use temperature: 3,630°F (2,000°C)
• Flexural strength at 3,000°F (1,650°C): 350 to 500 MPa
• Fracture toughness: 5 to 8 MPa·m^0.5
• Density: 5.3 to 5.9 g/cm³
• Thermal conductivity: 50 to 90 W/m·K (higher than most ceramics, aids thermal redistribution)

The condition under which UHTC oxidation protection works is temperature-dependent phase behavior. Below 2,730°F (1,500°C), the ZrB2 surface forms a porous ZrO2 scale with minimal B2O3 content, which does not self-seal and allows oxygen ingress. Above 2,730°F (1,500°C), sufficient B2O3 forms to create the protective liquid-phase-sealed oxide, and oxidation resistance improves markedly. This means UHTC components must reach operating temperature quickly or protected by an ablative outer layer during the initial heating phase.

The failure mode at the highest temperatures (above 3,600°F / 2,000°C) is active oxidation: at very low oxygen partial pressures combined with high temperatures, ZrO2 does not form a stable protective scale. Instead, volatile ZrO(g) forms and evaporates, causing rapid recession at rates of 0.1 to 0.5 mm per second in test arc jet conditions. Practical hypersonic vehicle design accounts for this by sizing the UHTC component with an ablative allowance: the leading edge is manufactured with extra material that ablates predictably during the highest-temperature phase of flight.

UHTC Applications: Scramjet Combustors and Hypersonic Glide Vehicles

Scramjet (supersonic combustion ramjet) combustors operate at Mach 5 to 15 with combustion chamber wall temperatures reaching 3,600°F to 5,400°F (2,000°C to 3,000°C) and pressures of 1 to 5 atm. Active cooling with cryogenic hydrogen fuel is the primary thermal management strategy, but the combustor liner material must survive the radiant and convective heat flux between fuel injection events.

NASA’s X-43A and X-51A Waverider programs demonstrated that ZrB2-SiC composite combustor panels survive short-duration scramjet operation (10 to 300 seconds). The panels sustain heat fluxes of 50 to 500 W/cm² without structural failure, compared to maximum nickel superalloy limits of approximately 10 to 20 W/cm² without active cooling.

AFRL’s High Mach Flight Research program documented that UHTC leading edge components on hypersonic glide vehicles must withstand not only high temperature but also high-velocity erosive particle impact from ablative debris. ZrB2-SiC composites with SiC fiber reinforcement show the best combined thermal and erosion resistance of all tested UHTC material systems as of current published research.

How Does Silicon Nitride Perform in Aerospace Engine Applications?

Silicon nitride (Si3N4) is used in aerospace bearing components, turbocharger rotors on military auxiliary power units (APU), and small turbine engine components where its combination of low density (3.2 g/cm³), high hardness (1,500 to 1,700 HV), thermal shock resistance, and corrosion resistance to jet fuel combustion products makes it superior to both steel and other ceramic options in specific applications.

Silicon nitride does not oxidize catastrophically below 2,550°F (1,400°C). It forms a protective SiO2 passivation layer on its surface when exposed to oxygen, and this layer, combined with the underlying Si3N4 matrix, resists further oxidation. This behavior, documented by Sheldon et al. in research published in the Journal of the American Ceramic Society, makes Si3N4 superior to SiC in oxidizing combustion environments at temperatures between 1,800°F (982°C) and 2,550°F (1,400°C) where SiC recession by water vapor attack is more problematic without an EBC.

Silicon Nitride Hybrid Bearings in Gas Turbine Engines

Hybrid bearings using Si3N4 balls in steel races are used in the main shaft bearings of gas turbine engines in the F-22 Raptor, F-35 Lightning II, and various commercial turbofan applications. The Si3N4 rolling elements provide three specific advantages over all-steel bearings in this application.

First, Si3N4 density is 3.2 g/cm³ versus 7.8 g/cm³ for bearing steel. A lower-mass rolling element produces lower centrifugal force at high shaft speeds (above 20,000 rpm), reducing race contact stress and extending bearing life by a factor of 3 to 5 in published fatigue testing at NASA Glenn Research Center.

Second, Si3N4 has a lower coefficient of thermal expansion (3.2 × 10⁻⁶/°C versus 12 × 10⁻⁶/°C for steel), so the rolling element does not expand as rapidly as the steel race during startup thermal transients. This reduces the risk of bearing seizure during cold-start operation at high altitude where lubricant viscosity is elevated.

Third, Si3N4 is electrically insulating (resistivity above 10¹³ Ω·cm), which prevents electrical discharge machining (EDM) damage when stray electrical currents, induced by variable-frequency drive systems or lightning strikes, pass through the bearing assembly. EDM pitting failure is a documented failure mode in all-steel aircraft generator drive bearings.

Key Specifications for aerospace-grade Si3N4 bearing elements (AS700 and NBD-700 grades from Saint-Gobain and CoorsTek):

• Density: 3.2 g/cm³
• Hardness: 1,500 to 1,700 HV
• Fracture toughness: 6 to 8 MPa·m^0.5
• Flexural strength: 700 to 900 MPa
• Maximum continuous operating temperature: 2,370°F (1,300°C)
• CTE: 3.0 to 3.2 × 10⁻⁶/°C

The failure mode for Si3N4 bearings is slow crack growth (SCG) from surface and subsurface flaws introduced during grinding of the bearing balls to ABEC precision tolerances. Si3N4 requires diamond grinding to achieve the spherical surface finish (Ra below 0.008 micrometers) required for bearing applications. If the grinding process introduces subsurface damage deeper than 5 micrometers, fatigue crack initiation occurs 10 to 100 times earlier than predicted by design models based on flaw-free material strength.

The fix is proof testing: every Si3N4 bearing ball for aerospace application is subjected to a proof load test at 1.5 to 2 times the maximum design contact stress. Balls that survive proof testing have demonstrated freedom from critical flaws, and their service life prediction becomes statistically reliable.

What Ceramic Coatings Protect Aerospace Structures Beyond the Engine?

Ceramic coatings in aerospace extend well beyond turbine engines and TPS tiles. Ceramic anodizing (hard anodize, Type III) protects aluminum airframe structures from fretting wear at fastener holes and sliding contact surfaces. Plasma electrolytic oxidation (PEO) converts the surface of titanium landing gear components into a ceramic oxide layer (TiO2 and Al2O3) 20 to 50 micrometers thick with hardness above 1,000 HV, far exceeding the 350 HV of the underlying titanium alloy.

Ceramic-filled PTFE dry film lubricants containing molybdenum disulfide (MoS2) and boron nitride (BN) particles are applied to flight control actuator sliding surfaces where liquid lubricants cannot be used at extreme altitude (below -65°F / -54°C) or would contaminate optical sensor windows.

Thermal Spray Ceramic Coatings for Abradable Seal Systems

Abradable seal coatings are applied to the inner surface of turbine and compressor shrouds. They are designed to be abraded by the blade tips as the engine heats up and the blade tips expand, creating a minimum-clearance seal that reduces air leakage around the blade tip.

Abradable seal coatings in the compressor section use Al2O3-TiO2 ceramic blends with 20 to 60% porosity, applied by air plasma spray (APS) to a thickness of 1 to 3 mm. The porosity allows the coating to abrade cleanly by the blade tip rather than loading the tip with hard ceramic debris that would cause tip wear or engine surge.

In the HPT section, abradable seal coatings use YSZ with 40 to 60% porosity (achieved by including hollow polymer spheres in the spray feedstock that burn out during coating deposition). HPT abradable seals must simultaneously function as a thermal barrier (low thermal conductivity) and an abradable material (low hardness at operating temperature), requiring a careful balance between YSZ phase content and porosity architecture.

According to a study published in Surface and Coatings Technology by Guo et al., optimizing HPT abradable seal clearance by 0.001 inch (0.025 mm) at the blade tip improves turbine aerodynamic efficiency by approximately 0.3 to 0.5%, a gain worth pursuing across thousands of flight cycles in commercial engine operation.

How Is Additive Manufacturing Changing Aerospace Ceramic Production?

Additive manufacturing (AM) of aerospace ceramics is moving from research demonstration to limited production certification. Three AM processes are currently most relevant to aerospace ceramic components: stereolithography (SLA) of ceramic-loaded photopolymer resins, binder jetting of ceramic powders, and directed energy deposition (DED) of ceramic coatings.

Ceramic SLA uses alumina, silicon carbide, or mullite particles suspended in a UV-curable resin at 40 to 65 vol% solids loading. The printed “green body” is then sintered in a furnace to burn out the polymer and densify the ceramic. Lithoz GmbH (Austria) and Admatec (Netherlands) produce commercial ceramic SLA systems capable of printing Al2O3 components with final density above 99% theoretical and surface roughness of Ra 1 to 3 micrometers without post-machining, adequate for many aerospace bracket and nozzle guide vane applications.

Binder Jetting of CMC Preforms

Binder jetting deposits a liquid binder onto layers of ceramic powder (SiC, Si3N4, or ZrB2) to build a 3D preform without requiring a die. The preform is then sintered or infiltrated with a ceramic or metal matrix phase. For SiC/SiC CMC, binder-jetted SiC preforms can be infiltrated with molten silicon (reaction bonded silicon carbide, RBSC) to create near-net-shape components with complex internal cooling channels that are impossible to machine into dense SiC.

NASA and GE Aviation have published collaborative research on binder-jetted SiC preforms for CMC turbine blade platforms. The AM preform approach reduces the number of woven fabric plies required (and therefore labor) for complex-curvature CMC components and enables internal channel geometries with hydraulic diameters below 1 mm, far smaller than any drilled cooling hole in a metal blade.

Key Specifications for binder-jetted RBSC components:

• Final density: 90 to 95% theoretical (versus 85 to 90% for CVI CMC)
• Surface roughness before post-processing: Ra 5 to 15 micrometers
• Minimum internal channel diameter achievable: 0.5 to 1.0 mm
• Flexural strength: 300 to 450 MPa (lower than hot-pressed SiC but adequate for shroud and platform applications)
• Dimensional tolerance post-sintering: ±0.3 to ±0.5% linear

The failure mode specific to binder-jetted ceramic preforms is binder burnout cracking. During the thermal debinding step (typically 400°F to 900°F / 200°C to 480°C), the binder polymer decomposes and leaves the preform temporarily without a bonding phase. If the heating rate exceeds 1°C to 2°C per minute during this phase, the vapor pressure of decomposing binder exceeds the green strength of the preform, causing macroscopic cracking that cannot be healed in subsequent sintering.

The ceramics knowledge required to understand AM ceramic densification overlaps significantly with traditional kiln firing science. For those interested in exploring how sintering and thermal processing relate to classical ceramic firing, our guide covering foundational ceramic processing concepts taught in ceramics education provides relevant background on thermal densification principles.

Aerospace Ceramics Compared: Material Properties Across Key Systems

Use the table below to compare maximum operating temperatures, density, thermal conductivity, fracture toughness, and primary aerospace application across the major ceramic material systems covered in this guide.

The weight advantage of all ceramic systems over nickel superalloy (8.2 g/cm³) is the first-order driver of adoption in rotating engine components, where weight reduction at high radius directly reduces the centrifugal loading the disk must carry.

How Do Ceramic Radomes Protect Radar and Missile Guidance Systems?

Radomes are ceramic or ceramic-composite structures that house radar antennas on aircraft nose sections and missile guidance heads. A radome must simultaneously survive aerodynamic heating, rain erosion at flight speed, and thermal shock while transmitting radar signals with less than 0.5 dB of electromagnetic signal loss across the operating frequency band.

Pyroceram (glass-ceramic) and fused silica (SiO2) radomes have been used in aircraft applications since the 1950s. Fused silica has a dielectric constant (εr) of 3.8 and a loss tangent of 0.0001, making it nearly transparent to radar across S, C, and X bands (2 to 12 GHz). Its CTE of 0.55 × 10⁻⁶/°C, the lowest of any bulk ceramic material, gives it exceptional thermal shock resistance: fused silica survives a thermal gradient of 1,800°F (1,000°C) in 1 second without cracking, documented by Corning in their Vycor 7900 glass-ceramic product data.

Hypersonic missile radomes face a more demanding combined environment. At Mach 5 (3,836 mph / 6,174 km/h), a missile nose cap experiences aerodynamic heating of 20 to 100 W/cm² depending on altitude and flight duration. Raytheon’s Standard Missile and AMRAAM use hot-pressed silicon nitride radome domes for supersonic applications, and development programs for Mach 5+ missiles are evaluating UHTC-based radomes with added dielectric control through compositional adjustment of the hafnium boride to hafnium oxide phase ratio.

The dielectric condition that constrains UHTC radome use is electrical conductivity. ZrB2 and HfB2 are electrically conductive (resistivity 10 to 60 μΩ·cm), making them opaque to radar. UHTC radome designs must either use the oxide phases (ZrO2, HfO2) which are dielectrically transparent but less thermally capable, or accept that the radome is ablative and functions only once. Current DARPA-funded programs target a ZrO2-stabilized hafnium oxide radome material that maintains εr below 10 and loss tangent below 0.01 while surviving up to 3,270°F (1,800°C) for the duration of a Mach 8 intercept trajectory.

Ceramics in Aerospace vs Defense: Where the Two Fields Intersect

Aerospace ceramics and defense ceramics share materials, manufacturing processes, and failure modes. Silicon carbide used in CMC turbine components is the same base material used in ceramic armor systems. Hot-pressing equipment that produces Si3N4 bearing balls for jet engines also produces Si3N4 plates evaluated for armor applications.

The key difference is the loading mode. Turbine component ceramics are optimized for sustained high-temperature mechanical loading under steady-state stress. Armor ceramics are optimized for ballistic impact resistance under instantaneous high-strain-rate loading. A turbine CMC with 20 to 25 MPa·m^0.5 fracture toughness designed for fatigue resistance is not automatically an effective armor material, because armor performance depends on hardness and the ceramic’s ability to blunt a projectile’s tip, which requires different microstructural optimization.

For readers interested in how the same ceramic materials used in engines and TPS perform under ballistic loading conditions, our analysis of how ceramic armor stops high-velocity projectiles explains the mechanics of ceramic defeat mechanisms using the same SiC and Al2O3 material families.

The commercial transfer runs in both directions. Fracture mechanics advances from armor research have improved the proof testing protocols used to qualify Si3N4 bearing balls for turbine applications. EBC coating development for CMC engine components drew on ceramic coating expertise developed initially for reactor shielding applications.

What Are the Current Limitations Preventing Wider Aerospace Ceramic Adoption?

The three barriers limiting broader aerospace ceramic deployment are cost, inspection difficulty, and the absence of ductile failure behavior. Each barrier is specific, quantifiable, and the subject of active industry programs.

Cost is the most cited barrier. CVI SiC/SiC CMC processing costs $1,000 to $3,000 per pound of finished component, compared to $200 to $600 per pound for nickel superalloy investment castings. The high cost reflects the 100 to 400-hour CVI deposition cycle time, the specialized high-temperature furnace infrastructure required, and the high scrap rate from porosity defects discovered during post-processing nondestructive testing (NDT). AM approaches (binder jetting, ceramic SLA) are actively targeting this cost barrier by reducing machining and preform labor, but certification of AM ceramic components requires additional rounds of qualification testing that delay the cost benefit realization.

Inspection difficulty is the second barrier. Monolithic ceramics and CMC components are opaque to ultrasonic inspection techniques optimized for metals. Ceramic CMC requires X-ray computed tomography (CT) at synchrotron radiation sources, industrial CT with microfocus X-ray tubes, or laser ultrasonic scanning to characterize internal fiber architecture, delamination, and void content. A single CMC turbine shroud requires 2 to 6 hours of CT scan time at a cost of $500 to $2,000 per component for production NDT, adding 10 to 20% to the finished part cost.

The absence of ductile failure behavior is the most fundamental limitation. Metal turbine components deform visibly before fracturing, giving maintenance crews a detectable warning. Ceramics fail by brittle fracture without visible deformation, and failure can be catastrophic and sudden. Engine certification authorities require demonstrated damage tolerance before approving ceramic components for life-limited rotating assemblies. The solution pursued by CMC developers is proof testing (every component is subjected to an overload before installation) combined with structural health monitoring (acoustic emission sensors embedded in the engine case that detect early CMC cracking) and periodic borescope inspection to look for matrix microcracking before it progresses to fiber fracture.

According to Bansal and Lamon in Ceramic Matrix Composites: Materials, Modeling and Technology, the probability of CMC component failure per flight hour can be brought below 10⁻⁹ (the industry standard for catastrophic failure rate per flight hour) only when all three of proof testing, NDT certification, and in-service health monitoring protocols are implemented together, not individually.

Frequently Asked Questions About Ceramics in Aerospace

Can CMC turbine components be repaired in the field, or do they require complete replacement?

Current SiC/SiC CMC turbine components cannot be field repaired and must be returned to an overhaul facility for any damage beyond surface coating wear. The CMC matrix repair process requires reinfiltration of SiC precursor into the crack network under high-temperature conditions (above 2,000°F / 1,093°C) that cannot be replicated in a maintenance hangar. Minor EBC coating spallation can be addressed by cold spray or low-temperature plasma spray at overhaul shops equipped with the coating systems. GE Aviation’s service data indicates HPT CMC shrouds have an on-wing life of 3,000 to 6,000 flight cycles before mandatory removal for inspection and potential coating refurbishment.

What is the difference between a thermal barrier coating and a thermal protection system?

A thermal barrier coating (TBC) is a thin ceramic layer (100 to 300 micrometers) bonded to a metal component to reduce metal surface temperature by 100°F to 300°F (55°C to 167°C). A thermal protection system (TPS) is a structural ceramic or ceramic composite system that replaces the metal structure entirely in the highest-heat zones. TBC operates on nickel superalloy substrates inside operating turbine engines at sustained temperatures for thousands of hours. TPS operates on spacecraft structures during atmospheric reentry for minutes to tens of minutes. The two systems share ceramic base materials (YSZ, silica, alumina) but differ in thickness, structural role, loading duration, and thermal cycling type.

Why is YSZ used for thermal barrier coatings instead of pure zirconia?

Pure zirconia (ZrO2) undergoes a destructive phase transformation from tetragonal to monoclinic at 2,120°F (1,160°C) on cooling, accompanied by a 3 to 5% volume expansion that cracks any coating within a few thermal cycles. Adding 7 to 8 wt% yttria (Y2O3) stabilizes the tetragonal phase and prevents this transformation down to room temperature. The stabilized tetragonal phase also has lower thermal conductivity (1.5 to 2.0 W/m·K versus 2.5 to 3.0 W/m·K for cubic-stabilized ZrO2) due to increased oxygen vacancy concentration. The 7 to 8 wt% yttria content is a specific optimum: lower yttria content does not fully stabilize the phase, and higher content produces fully cubic YSZ with higher thermal conductivity and lower fracture toughness.

Do aerospace ceramics ever appear in structural airframe components, or only in thermal and engine applications?

Aerospace ceramics appear in selected structural airframe roles beyond thermal and engine applications. Alumina ceramic radomes on fighter aircraft nose sections are load-bearing structures that must withstand aerodynamic pressure and bird strike loading. Boron carbide (B4C) ceramic tiles are used as ballistic armor inserts in military helicopter crew seats and floor panels. Ceramic bearings in flight control actuators (Si3N4 balls in titanium races) handle structural loads in primary flight control linkages on the F-35. Fused silica ceramic windows on cockpit canopies resist laser dazzle weapons by absorbing specific UV and near-UV wavelengths. These applications are smaller in volume than engine ceramics but represent a broader role for advanced ceramics in total aircraft systems.

How does water vapor affect SiC/SiC CMC, and why does the engine environment make this worse than a furnace environment?

Water vapor causes SiC to oxidize in a two-step reaction: SiC first forms a SiO2 scale (passive oxidation), then water vapor reacts with the SiO2 to form volatile Si(OH)4 gas (active volatilization), removing the protective scale and exposing fresh SiC. In a laboratory furnace at atmospheric pressure, passive oxidation dominates and the SiO2 scale is stable. In a gas turbine combustor, water vapor partial pressure is 5 to 15% of total pressure (versus less than 1% in ambient air), and the high gas velocity (50 to 100 m/s) continuously removes the volatile Si(OH)4 product before it can re-deposit, maintaining the driving force for SiC recession. At typical HPT conditions (2,100°F / 1,150°C, 10% H2O, 100 m/s gas flow), SiC/SiC CMC without an EBC recedes at 0.001 to 0.01 mm per 1,000 hours of engine operation, making a 20-year engine service life without EBC impractical.

Are aerospace ceramics food safe if repurposed in kitchen applications?

Aerospace ceramic materials such as alumina, silicon nitride, and zirconia are chemically inert and non-toxic in bulk form, and alumina is widely used in food-contact applications including cutting inserts and conveyor components in food processing equipment. However, aerospace-grade CMC components contain silicon carbide fibers that can release fine ceramic particles if the component is cut or abraded, and these should not be used in food contact roles. YSZ thermal barrier coatings on turbine blades contain small quantities of yttrium oxide, which has low acute toxicity but whose chronic dietary exposure effects are not established for human consumption. There is no practical or economic scenario in which aerospace ceramic components would be repurposed for food contact use, and no food-safety certifications exist for aerospace ceramic grades.

What ceramic materials are used in solid rocket motor nozzles, and how do they differ from turbine ceramics?

Solid rocket motor (SRM) nozzle throats and exit cones use carbon-carbon composites, graphite-phenolic ablatives, and tungsten-reinforced carbon composites rather than the oxide or SiC ceramics used in turbine engines. SRM nozzles operate at 5,400°F to 7,200°F (3,000°C to 4,000°C) for 10 to 200 seconds, a duration too short to justify the cost of reusable ceramic systems. The SRM combustion environment is highly reducing (oxygen-deficient), which makes ablative carbon systems effective because they ablate predictably and carbon oxidation rates are low without free oxygen. Turbine ceramics (CMC, YSZ TBC) operate in highly oxidizing environments at lower temperatures for thousands of hours, requiring oxidation-resistant materials rather than ablation-tolerant ones. The two application classes use different material families for this reason, even though peak temperatures in SRM nozzles far exceed turbine engine temperatures.

How do ceramic components in aerospace get certified for flight under FAA or military airworthiness standards?

FAA certification of ceramic components on commercial aircraft follows Advisory Circular AC 33.15 for turbine engine materials, which requires a full material characterization database (B-basis and A-basis allowables from statistically significant test populations), a damage tolerance analysis demonstrating that the component survives a defined set of discrete damage events without catastrophic failure, and validation by engine cyclic endurance testing per 14 CFR 33.93. For CMC specifically, FAA issued Special Conditions for the GE9X engine covering CMC HPT shrouds and CMC combustor liners, requiring proof testing of every flight-critical CMC part and a demonstrated recession model validated against 5,000-hour engine testing. Military airworthiness for ceramic components follows MIL-HDBK-17-4 (ceramic matrix composite structural analysis) and MIL-STD-810 (environmental testing), with additional program-specific qualification requirements defined in the applicable MIL-SPEC for each platform.

Can silicon carbide foam replace solid SiC in aerospace heat exchangers?

Open-cell silicon carbide foam (10 to 45 pores per inch, 5 to 15% relative density) is used in aerospace heat exchanger applications where the goal is to maximize surface area for convective heat transfer within a minimum volume. SiC foam heat exchangers appear in satellite thermal control systems, hypersonic vehicle fuel-cooling circuits, and research scramjet pre-cooler designs. SiC foam provides specific surface areas of 1,000 to 3,000 m²/m³ at pressure drops of 10 to 100 Pa/cm, significantly higher surface area per unit pressure drop than conventional tube-and-fin heat exchangers. The limitation is compressive strength: SiC foam at 10% relative density has compressive strength of only 1 to 5 MPa, restricting its use to low-pressure environments or applications where the foam is encased in a structural housing that carries all mechanical loads.

What is the role of boron nitride in aerospace ceramics?

Hexagonal boron nitride (h-BN) serves two distinct roles in aerospace ceramics. As a fiber interphase coating in CMC, a 0.1 to 0.5 micrometer layer of h-BN is deposited on SiC fibers before matrix infiltration by chemical vapor deposition. This h-BN interphase is weak in shear (shear strength 10 to 30 MPa) and promotes crack deflection along the fiber-matrix interface rather than through the fiber, which is the mechanism responsible for CMC toughening. Without the h-BN interphase, SiC fibers bond too strongly to the SiC matrix, and the composite fails in a brittle manner indistinguishable from monolithic SiC. As a machineable structural ceramic, hot-pressed BN tiles and rods are used in plasma-facing components of satellite electric thrusters (Hall-effect and ion thruster grids) because BN has zero electrical conductivity, good thermal shock resistance, and is easily machinable to tight tolerances without diamond tooling.

How does the thermal conductivity of aerospace ceramics compare to the clay bodies used in traditional pottery?

Traditional stoneware clay bodies fired to cone 10 have thermal conductivity of 1.0 to 2.0 W/m·K. Porcelain fired to cone 10 measures 1.5 to 2.5 W/m·K. These values are similar to or slightly higher than YSZ TBC (1.5 to 2.0 W/m·K) but far below SiC/SiC CMC (18 to 25 W/m·K) and ZrB2 UHTC (50 to 90 W/m·K). LI-900 TPS tile (0.02 to 0.06 W/m·K) is 20 to 100 times more insulating than traditional ceramics, reflecting its 94% void fraction architecture. The comparison illustrates that thermal conductivity in ceramics spans three orders of magnitude depending on density, phase composition, and microstructure, and the aerospace application determines whether high conductivity (heat spreading), low conductivity (insulation), or intermediate conductivity (combined structural and thermal performance) is the design target.

What safety precautions are required when handling aerospace ceramic fibers and powders?

SiC fibers used in CMC manufacturing are classified as man-made vitreous fibers (MMVF) and present a respiratory hazard if inhaled as respirable-size particles (below 10 micrometers aerodynamic diameter). OSHA recommends engineering controls (enclosed processing, local exhaust ventilation) and a half-face respirator rated for particulates (minimum N95, NIOSH-certified) when cutting, grinding, or otherwise disturbing SiC fiber preforms or CMC components. Zirconia powder (for TBC feedstock preparation) and hafnium boride powder (for UHTC processing) require the same respiratory precautions and additionally require explosion-proof handling equipment because fine reactive metal boride powders are combustible at elevated temperatures. YSZ plasma spray operations generate an aerosol of molten ceramic particles and require full-face shields and fire-resistant protective clothing. None of these materials require specialized chemical hazardous waste disposal under RCRA unless contaminated with heavy metal coatings; clean SiC and Al2O3 grinding debris is classified as non-hazardous solid waste in the United States under current EPA guidelines.

Aerospace ceramic processing shares some dust and particulate hazard characteristics with traditional ceramics studio practice. Anyone working with fine ceramic powders in any application should review respiratory safety principles, and our overview of studio safety fundamentals including silica dust management covers foundational protective practices applicable across all ceramic powder environments.

The Future of Aerospace Ceramics: Materials on the Horizon

Three material systems currently in advanced research represent the most significant near-term additions to the aerospace ceramic toolbox: high-entropy ceramics (HEC), silicon carbide with hexagonal boron nitride nanoplatelet reinforcement, and thermally stable rare-earth monosilicate EBC systems.

High-entropy ceramics contain five or more principal cation species in equimolar or near-equimolar ratios on a shared crystal structure. (Hf,Zr,Ti,Ce,Y)O2 high-entropy oxide ceramics demonstrated at Oak Ridge National Laboratory show thermal conductivity below 1.0 W/m·K at 1,832°F (1,000°C), lower than any current single-composition TBC material, combined with phase stability to at least 2,730°F (1,500°C). If validated in cyclic thermal fatigue testing, high-entropy oxide TBC could extend HPT blade ceramic coating life by 50% compared to 7 wt% YSZ while operating 300°F (167°C) higher.

SiC reinforced with hexagonal boron nitride nanoplatelets (h-BN-np) at 2 to 5 vol% loading improves fracture toughness from 3 MPa·m^0.5 for monolithic SiC to 6 to 8 MPa·m^0.5 by crack deflection along the low-shear-strength h-BN basal planes, without sacrificing the hardness or oxidation resistance of the SiC matrix. This improvement makes monolithic SiC-BN-np a potential replacement for the more expensive and process-intensive SiC/SiC CMC in applications below 1,800°F (982°C) where full CMC fiber architecture is not required.

Ytterbium disilicate (Yb2Si2O7) EBC topcoats are replacing BSAS-based EBC systems in next-generation CMC engine programs because Yb2Si2O7 has lower water vapor reactivity than BSAS, a CTE closer to SiC (4.7 × 10⁻⁶/°C versus 5.0 × 10⁻⁶/°C for SiC), and phase stability to 2,550°F (1,400°C). GE Aerospace and Rolls-Royce have both published EBC development roadmaps that include Yb2Si2O7 as the next-generation topcoat for CMC turbine airfoils in engines targeting turbine inlet temperatures above 3,000°F (1,650°C).

The progression from traditional ceramics to these advanced aerospace material systems represents a direct extension of the same silicate and oxide chemistry that potters and ceramic artists have worked with for centuries, scaled to atomic-level microstructural engineering and validated under conditions no kiln can replicate. For a comprehensive view of how traditional ceramic materials relate to these advanced technical ceramic families, our reference on the full spectrum of ceramic material categories maps every class from earthenware to aerospace-grade composites in a single comparative framework.

Aerospace ceramics will continue to expand into rotating turbine components, hypersonic structures, and propulsion system liners as CMC manufacturing costs fall and AM ceramic qualification frameworks mature. The discipline of understanding what ceramics can and cannot do at the material science level, which starts in the studio and extends to the engine test cell, remains the same across every application.

Here is the comparison of aerospace ceramic material systems for thermal protection and engine applications across all key performance dimensions.

Aerospace ceramics represent the highest-performance end of a materials family that spans from studio earthenware to hypersonic leading edge composites. Understanding the mechanisms behind thermal resistance, the conditions under which each ceramic system functions or fails, and the specific engineering trade-offs between cost, weight, and temperature capability gives any reader of this guide a working framework for evaluating aerospace ceramic technologies as they move from research to certified production in the next generation of aircraft and spacecraft.