Material selection for Mechanical Design (ME 770)

TERM PAPER

Material Selection For High Temperature Aerospace Application(Nose Cone)

Submitted to: Submitted by:

Prof. Kamal K Kar Shree Ram Pandey

Professor ,Mechanical Engg. Department (14105290)

ABSTRACT

Given the problem of the aerodynamic design of the nose cone section of any vehicle or bodymeant to travel through a compressible fluid medium (such as a rocketor aircraft, missile or bullet), an important problem is the determination of the nose conematerial for safety and optimum performance. For many applications, such a task requiresthe recognition of several involved process parameters like heat flux, drag acting,temperature gradients etc. Materials have to be tested for several properties like thermalinsulation, thermal shock, thermal distortion etc. This paper makes an attempt to addresssome of the several issues arising during material selection of nose cones in aerospaceapplication.

Keywords : Nose cones, High temperature strength, Temperature gradient,

Ceramics,

INTRODUCTION:

During a launch vehicle ascent through denser atmosphere the vehicle is subjected to thermalloads on its exterior due to aerodynamic heating. These thermal loads manifest themselves inthe increased skin temperature for the structure. The skin section which attain hightemperatures becomes the heat source for the various electronic packages and controlcomponents housed within various housings such as nose cone fairing and interstagecompartments. Thus, there is a need to provide thermal protection to the structure to ensureenvironmental protection for various packages and structure temperature to within usefultemperature range for the material. The component of the aerospace structure which is mostseverly affected is nose cone.

NOSE CONE:

The nose cone is the forward most section of the guided missile. Due to high temperaturesexperienced at the leading edge of the nose cones in case of high speed applications, it isgenerally made of materials which can sustain high temperature and the resultant expansionload due to that heating.

On a rocket vehicle it consists of a chamber or chambers in which a satellite, instruments,animals, plants, or auxiliary equipment may be carried, and an outer surface built towithstand high temperatures generated by aerodynamic heating. Much of the fundamentalresearch related to hypersonic flight was done towards creating viable nose cone designs forthe atmospheric re-entry of spacecraft and ICBM re-entry vehicles.

In a satellite vehicle, the nose cone may become the satellite itself after separating from thefinal stage of the rocket or it may be used to shield the satellite until orbital speed isaccomplished, then separating from the satellite.

On airliners the nose cone is also a radome protecting the weather radar from aerodynamicforces.

The shape of the nose cone must be chosen for minimum drag so a solid of revolution is usedthat gives least resistance to motion.

CURRENT ISSUES:

In the recent years space vehicles like rockets, re-entry vehicles regardless their uniquedesigns needed control surfaces at hypersonic speeds. Low-radius leading edges are subjectto much greater aerodynamic heating than blunt edges, such as those on the Space Shuttle,and they thus will reach temperatures that may exceed 2000C during re-entry. Availablethermal protection materials will not survive such extreme temperatures and new materialsare required for advanced thermal protection systems. Thus they need advanced and latestultra high temperature resistant materials [2006] used as a Thermal protection layer(TPS)layers which have high resistance in heat and oxidation.

Challenges posed in material selection for nose cone:

High heat flux over small areaHigh temperature, oxidation, erosionVery high temperature gradients

Materials Selection Procedure:

Terminology:

Concept design: Basic design of a product to meet the main functionalobjectives and performance requirements determined from market research.A large number of materials are considered at the concept design stage.

Constraint (design): Performance requirements that must be met by theproduct (and its materials).

Detailed design: Final stage of the design process involving all the detaileddesign work to complete the product. Also involves converting the designinto specifications and documentation so the product can be produced.The material to be used in the product is selected in the detailed designstage.

Embodiment design: Determination of shape, size and other major designfeatures of the product. A shortlist of candidate materials for the productare considered in the embodiment design stage.

Function (design): The purpose of the designed product.

Material indices: Quantitative measure of how well a material property (e.g.stiffness, strength, maximum operating temperature) exceeds the designconstraint. Index values are used to rank shortlisted materials in order ofexcellence to exceed the design constraint limit.

Objective (design): The aims of the design, such as to minimise cost ormaximise fatigue life.

Ranking: Listing shortlisted materials in their order of excellence to exceedthe design constraint limit based on the index value.

Screening: Process of eliminating materials based on their inability to meetone or more of the design constraints.Supporting information: Various resources (e.g. case histories, past experiences)used in the final stage of the materials selection process.

Translation: Process of translating the design into functions and objectiveswhich can used to define the constraints for materials selection.

Table 1 Design-limiting properties used in materials selection

Class Property

Structural DensityElastic modulus (tension, shear, etc.)Strength (yield, ultimate, fracture)Impact damage resistanceHardnessFracture toughness (damage tolerance)Fatigue (life, strength)Creep resistance (creep rate, stress rupture life)

Economic and business Cost (raw material, processing, maintenance)AvailabilityService lifeRegulatory issues

Manufacturing Fabrication and casting (formability, machinability,welding)Dimensional (shape, surface finish, tolerances, flatness)Number of itemsNondestructive inspection for quality assurance

Environmental durability Corrosion rateOxidation rateMoisture absorption rate

Environmental impact SustainabilityGreenhouse and other emissions during manufactureRecyclingWaste disposal (hazardous)Health hazards (carcinogens, flammable)

Specialist Thermal conductivityElectrical conductivityThermal expansionThermal shock resistanceStealth (electromagnetic absorbance/infrared)

STAGES OF MATERIAL SELECTION:

Fig: The four steps of material selection

Material Screening:

Screening involves eliminating those materials whose properties do not meet the designconstraints. The constraint defines an absolute upper or lower limit on property values, andmaterials that do not meet the limiting value are screened out. No trade-off beyond this limitis allowed. In addition to constraints on the mechanical properties, other constraints may beapplied related to economic/business, manufacturing and environmental factors as well asspecialist properties. These can also be used to screen out materials.

Material Indices:

A material index measures how well a candidate material that has passed the screening stepcan do the job required by the component. There are many material indices, each associatedwith maximising or minimising some property value, such as maximum strength per unitweight or minimum manufacturing cost per unit product.

Some of the material indices for the nose cone problem can be mentioned in the table below:

Design objective Material Index

Maximise thermal insulation 1/k

Minimise thermal distortion k/a

Thermal shock resistance

Maximise damage tolerance (beam, plate, etc.) Kic

Maximum pressure vessel strength Kic

E: Young's modulus;

k: thermal conductivity;

: thermal expansion coefficient;

f: failure strength;

Kic: fracture toughness.

Equations such as these are used to calculate the index value for each candidate material thatpasses the screening stage, and then the materials are ranked in order of excellence. It is alsoimportant to consider at this stage whether the materials can be fabricated into thecomponent. There is no point ranking a material if it cannot be cost-effectively processed intothe final product.

Supporting information and final selection

Once the shortlist of candidate materials are ranked in order of excellence using the indexvalues, the final stage of the selection process is performed, and this involves the use ofsupporting information for a detailed profile of each material. Supporting informationinvolves important factors other than the material properties that are relevant to the design,such as previous uses of the material in similar applications; the availability of the material;whether the company has prior manufacturing experience with the materials; certificationissues associated with the material (i.e. has it been previously certified by aviationregulators); whether the material has any special handling requirements or poses occupationalhealth and safety problems during manufacturing; whether the material can be recycled; andso on. Many sources, including databases and case histories, are used to collect as muchinformation as possible about each material. The supporting information is analysed for eachcandidate material, and from this the final material is selected.

The process of screening and ranking materials on their properties can be exhaustive when anextremely large number of materials are under consideration. With over 120 000 materialsbeing available, the task of individually assessing each material against the objectives andconstraints of the design is not practical. Material property charts are used to rapidlyscreen out large numbers of materials and to identify those materials that meet the propertyconstraint. Material property charts plot two properties, as shown below:

Fig. Material chart between Linear expansion coefficient ad thermal conductivity

Fig. Material chart between strength and temperature(representational only)

These charts condense a large body of property data into a compact yet accessible and easy tounderstand form. The charts also reveal correlations between material properties, such as therelationship between strength and toughness or between elastic modulusand density capacity.

Fig. Temperature distribution for sharp and blunt edges

Various Approaches to the modelling:

Real time modelling:

The most accurate way to find the aerodynamic heating on a high velocity flight structure isto fly it and measure what happens. This approach can be expensive, though, requiringbuilding

Computational modelling:

Another approach is to model the rocket in Computational Fluid Dynamics (CFD) software.This works well for steady state processes, but is incredibly processor-intensive for a time-dependent process such as trying to find the total heating of a nose cone in a flight trajectorywith continuously changing air velocity, pressure, and temperature. A steady state analysiscan be done with high fidelity results in under an hour, but a 40 second flight simulationrequires about a month of processor time.

Empirical modelling:

Rockets and planes have been flown supersonic many times, and from that datasetresearchers have created simple equations that can crudely predict the temperature andheating that will be seen in flight. This is least accurate.

Recovery Temperature

Stagnation temperature is the highest temperature of the gas possible, converting all of itskinetic energy into heat.

stagnation temperature but reflects the fact that the conversion to heat from dynamic pressure

TRecovery = TFree Stream ( 1 + 0.2 r M2 )

Where:

TRecovery is the recovery temperature, in Kelvin or Rankine.

TFree Streamtemperature at the given altitude, in Kelvin or Rankine.

r is the recovery factor. For stagnation r = 1; turbulent boundary layer r = 0.9;

M is the mach number.

Altitude(m)

Temp(K)

Pressure(Pa)

Density(kg/m^3)

Dynamicpress(Pa)

Machnumber

Stagn.temp(K)

Recoverytemp (K)

827 282.8 91777 1.131 66898 1.02 341.7 335.8

2022 275.0 79278 1.004 142116 1.60 415.9 401.8

8056 235.8 35312 0.522 33808 1.17 300.3 293.8

9955 223.4 26619 0.415 126247 2.60 526.2 495.9

11750 216.7 20108 0.323 199903 3.77 831.9 770.4

13703 216.7 14771 0.238 247967 4.90 1255.6 1151.7

TABLE 2 Variation of stagnation temperature and recovery temperature with altitude

stagnation temperature of the air. It gives an upper bound to thetemperature that the nose could reach.

As we can see, the recovery temperature is a bit lower than the stagnation temperature, but

Analysis of the Nose Cone

To analyze the heat going(essentially thermal load analysis) into the nose cone its materialproperties must be defined. These include:

Density

Specific heat (possibly at multiple temperatures)

Conductivity type (Isotropic, Unidirectional, Asymmetrical/Biaxial, or Orthotropic)

Thermal conductivity (which can also vary with temp and can be multiple, if notisotropic)

Melting temperature (the glass transition temperature of the resin would be a goodanalog for a composite nose cone)

Thermal Analysis

Thermal analysis for the nose cone during ascent proved the limiting factor throughout thedesign phase. An initial analysis of the power-law body as originally defined immediatelyproves that the heating rate at the tip of the nose would approach infinity, implying infiniteheat transfer to the nose cone throughout flight. As an infinite heating rate is clearlyunacceptable, the first step required blunting the tip of the nose cone in order to bring theradius of curvature up. The heating rate of a leading body is dependent upon both thephysical shape of the object as well as the material properties. Heating rate is primarilydependent upon the radius of curvature of the test body at a specific point as well as thespecific heat of the material used. The heating rate of a leading edge body can betheoretically determined using Eq. (3) below.

where ,q is the heating rate per unit area,

is the density of the fluid,

rn is the radius of curvature of test body,

V is the instantaneous velocity, cpw is the specific heat of surface material, and

Tw is the instantaneous temperature at surface.

We can see from Eq. (3) that the heating rate is dependent upon trajectory, material andstructural parameters. Since our design process do not entail changing the optimal trajectoryand therefore the velocity at any point in the launch, we are forced to focus on changes toboth the material and structural properties. Ideal design for meeting the thermal requirementswould entail increasing the radius of curvature throughout the nose cone, especially at thestagnation point, as well as employing a material with a higher specific heat. Eq (3) clearlyshows that as the radius of curvature at a point decreases, it increases the instantaneous heattransfer, which accumulates throughout the flight. Qualitative analysis alone is able to provethat the original power-law body was unsuited to withstanding high velocity flight, whichrequired using a simplified thermal analysis model with a blunted tip.

Fig. Meshing(Obtained in SOLIDWORKS)

The initial heating rate equation requires a complicated iterative process as well as convertingthe given heating rate from Eq. (3) to a heating rate per volume and then an overalltemperature. Initial steps to determine this heating rate required a calculation of both thelocal atmospheric enthalpy as well as the velocity contribution. The local, atmosphericenthalpy is calculated using Eq. (4) below:

where ha is the local, atmospheric enthalpy, Cp is the specific heat of air, defined as 1003.5kJ/kg-K and T is the temperature at the desired altitude calculated using StandardAtmosphere tables.

The velocity contribution is the 0.5V2 term, which contributes more to the conditions on thesurface of the nose cone due to our high velocity through high altitude/low-density

combined value. This allows us to determine the local conditions that will have an effect onthe heating rate of the nose cone. Figure 3 shows that since we are launching from a balloonat approximately 30km, the local atmospheric enthalpy contributes very little to the overallenthalpy. As expected with a squared term, the velocity contribution increases slowly at firstand then rapidly as the velocity continues to increase throughout ascent.

Heat flux

Tactical Missile Design also provides a very imperial empirical equation for calculating theheat flux the nose cone will see.

0.8 M2.8/x0.2

Where:

Q is the heat flux in BTU per square foot per second.

M is the mach number.

x is the distance from the tip, in feet.

With this equation we get the following chart for the heating of the nose at various positionsat the high velocity data point:

Fig. Variation of Recovery temperature on Mach Number

Fig. Temperature variation over surface (ANSYS)The shockwave coming off the cone is prominent in all the plots, and raked back at a severeangle due to the high velocity. The temperature is relatively low, aside from a small warmarea just downstream of the nose.Since the rockets and airships move with a high velocity in the compressible medium,itbecomes necessary to study the variation of gas density with velocity and temperature.

Fig. Gas Density(ANSYS)

Structural Analysis

Once the nose is capable of handling the thermal loading expected during ascent, we begin toanalyze the structural properties of the nose cone and how well it would survive the physicalloading due to ascent. Of primary concern in this analysis is the stagnation pressure on theblunt nose during ascent. Similar to the method used to determine the total enthalpy duringthe ascent, local atmospheric pressure was calculated as a function of time during the ascentusing the Standard Atmospheric Tables while dynamic pressure was calculated using theabsolute velocity data provided by the Trajectory group. Stagnation pressure was thereforecalculated using Eq. (5) below:

Where

Ps is the desired stagnation pressure,

Pa is the local atmospheric pressure from the Standard Atmosphere tables,

is the density of air at the current altitude

V is the absolute velocity of the launch vehicle.

Fig. Pressure distribution over the surface

High Temperature Application Materials:

High-Temperature Metals:

In general, metals have much greater ductility and higher densities than either ceramic ornonmetal matrix composites. This means that they are not subject to brittle failure and at thesame time have lower specific strength and stiffness. Intrinsically, most metals have higherthermal and electrical conductivities than either ceramics or nonmetal matrix composites(except carbon carbon composites). The three distinct types of high-temperature metals aresuperalloys, platinum group metals, and refractory metals.

Superalloys are usually based on group VIIIA elements (Ni, Fe, and Co). They have beendeveloped for service up to 1100C where they must maintain excellent mechanical strength,phase stability, resistance to creep, high surface stability, and resistance to corrosion andoxidation under relatively severe mechanical stresses. Examples of well-known superalloyfamilies are Hastelloy, Haynes, Inconel, Rene, Monel, Incoloy, and Waspaloy. These alloysare utilized in the hot sections of industrial gas turbines, turbocharger turbines, and marineturbines, as well as for turbine blades in the hot section of jet engines (Figure 3.1).The platinum group metals (platinum, rhodium, and iridium) with melting points rangingfrom 1770C to 2450C function well under mechanical loads and simultaneous corrosiveattack since they are chemically stable, as well as being resistant to oxidation and reactionwith many molten oxides. Most people know of the platinum group metals and their alloys asthermocouples, but in fact they are indispensable in many areas requiring high temperaturesand resistance to chemical attack. Applications include glass melting and spinningequipment, small rocket nozzles for station-keeping, single crystal growing, as well ascapsules for radioactive power sources.

The refractory metals: tantalum, niobium, tungsten, molybdenum, and rhenium all havemelting points in excess of 2400C (2477 3400C), are extraordinarily resistant to heat,wear, and creep as well as being relatively chemically inert. However, they are not aschemically stable and oxidation resistant as the platinum group. This is why iridium is coatedon top of the more ductile rhenium as an oxidation protection coating above 2300C forcarbon carbon composites. These metals are also used, for example, in forging jet engines,rocket nozzles, and incandescent lamp filaments.

Ceramics:

By nature of their composition, virtually all ceramic materials are high-temperature materialsand some are ultra-high-temperature materials. Ceramics encompass a very broad spectrumof refractory materials that makes a universal definition difficult. In general, ceramicmaterials are inorganic, nonmetallic materials made from compounds of a metal and anonmetal. Ceramic materials may be crystalline, partly crystalline, or amorphous (i.e.,glasses) and the vast majority are held together with ionic or covalent bonds. However itshould be noted that a few ceramics (e.g., TaC) are held together by metallic bonds. Incontrast to metals, the strong bonding in ceramics causes these materials to fracture before

they are able to plastically deform if the temperature is below the ductile-to-brittle transition(DBT) temperature.

Ceramics enjoy numerous high-temperature applications because they offer a variety ofattractive properties including high stiffness, high strength in compression, great thermalstability, low thermal expansion, and extremely high melting temperature. In addition, theyare resistant to most forms of chemical (corrosion, oxidation), and physical (abrasion, wear)attack. Applications include furnace elements and insulation, catalytic converters , ballbearings, thermal barriers, fuel cells, and fuels for nuclear reactors.

Ultra-high-temperature ceramics (UHTC) are a class of ceramic materials with melting pointsin excess of 3000C that are used in environments that require strength as well asenvironmental resistance to erosion and ablation at extremely high temperatures.Applications for these materials include rocket propulsion, scramjet propulsion, andhypersonic flight either as monolithic structures, composites, or coatings.

Ceramics that meet these criteria are principally carbides and borides of titanium, zirconium,niobium, tantalum, and hafnium. The carbides tend to have lower oxidation resistance thanborides at intermediate temperatures due to the formation of CO gas as one of the oxidationproducts. However, they have higher melting points (at 3997C, TaC has the highest meltingtemperature of any material) and perform well in erosive environments such as rocketnozzles. Borides on the other hand, have higher thermal conductivity than many ceramics(60 and are used in applications where heat needs to be moved and spread,such as the leading edges of hypersonic vehicles.

High Temperature Composites:

Composites comprise a large family of materials in which a reinforcement phase is placed ina continuous matrix phase resulting in properties that are not possible to obtain with a singlematerial. The reinforcement can be in the form of a particulate, whisker, continuous ordiscontinuous fiber, nanotube or nano-phase particle. In high-temperature composites thereinforcement and the matrix material are limited to ceramics, metals, and carbon. Thepossibility of combining these various material systems results in almost unlimited variationin composition and properties because the composites are tailorable. Due to the fact thatcomposites are more complex materials and less well known than metals and ceramics, theywill be covered in greater detail.

The possible variations in composites, and thus their properties, are further increased by thefact that even after selecting the composition of the matrix and the reinforcement, as well asthe type of reinforcement, there are many possibilities in how the reinforcement is placed inthe composite. That is, for example, the reinforcement can be dispersed in the matrix(particles, whiskers, discontinuous fibers, etc.) leading to isotropic properties, or in the caseof fibers (both continuous or discontinuous) a preform can be fabricated and subsequentlyfilled with a matrix resulting in anisotropic properties. Preforms are fabricated to the desiredshape by a variety different means. For example, 1-D preforms as well as 2-D fabric lay-up

preforms can be fabricated by hand, whereas braided, 3-D to 4-D (and even up to 11-D)woven preforms, as well as felted preforms require complex looms and other machinery.

It should also be noted that the enhanced properties of composites also result from thefunction of their constituents. That is, the matrix and the reinforcement play complementaryroles. For example, in a fiber-reinforced composite the fibers are the principal load-bearingcomponent. The matrix binds the fibers together, holding them aligned to carry the load, andtransfers the load applied to the composite through the fiber matrix bond to the fibersenabling the composite to withstand compression, flexural and shear forces as well as tensileloads. In addition, the matrix isolates the fibers from one another so that cracks are unable topass through all the fibers at one time which would result in brittle failure.

When pairing together a matrix material with a fiber, it is important that there be thermalcompatibility and chemical compatibility. High processing or use temperatures can lead tomatrix (or fiber) cracking during cooling when a thermal expansion mismatch is present.Chemical compatibility also prevents degradation at the fiber matrix interface at elevatedprocessing, heat-treating, and use temperatures. Degradation can be caused by chemicalreactions between the materials or phase changes in either component.

Coatings are often applied to protect the fibers from chemical attack, such as when usingcarbon fibers in a silicon carbide matrix. Fiber coatings are also employed to tailor theinterfacial bond strength, that is, the strength of the bond between the fiber and the matrix.This bond strength has a great effect on composite properties, such as decreasing toughnesswhen it is too strong.

Carbon-Carbon Composites:

Carbon carbon (C C) composites are a unique class of high-temperature and ultra-high-temperature materials due to their extraordinary combination of properties. These materialsare stronger and stiffer than steel as well as having a lower density (1.5aluminum. Their thermal shock resistance, and their excellent specific properties coupledwith the fact that their mechanical properties actually increase with temperature (>3200C)make them unsurpassed as ultra-high temperature structural materials.

Owing to their high performance and relatively high cost, C-C composites are used prin-cipally in the aerospace and astronautics industries. These materials are used to fabricaterocket nozzles, nose tips, exit cones, leading edges, and engine inlets of hypersonic vehicles(Figure 3.3) as well as high-temperature insulation, furnace hardware (Figure 3.4), andheaters for high-temperature applications. For reuseable hypersonic vehicles, the fact thatcarbon does not go through phase changes like some ceramics makes it a very valuablematerial. For satellite applications and high power laser mirrors, the high specific strengthand stiffness of carbon carbon composites as well as their near zero thermal expansionmakes them an ideal material for structures that require dimensional stability as they circlethe Earth.

The highest volume application of carbon carbon composites is as stators and rotors in thebraking systems of all military aircraft and most civilian jet aircraft (Figure 3.5). Thesecomposites enjoy a monopoly in this field because, in addition to the above mentionedproperties, they also possess high thermal conductivity, good frictional properties, and lowwear. This one application alone in which temperatures can reach 1400C on a rejected takeoff makes them the highest volume high-temperature composite

C C composites are fabricated through a multi-step process. Being a broad class of materials,the thermal and mechanical properties of carbon carbon composites vary greatly dependingon the type and grade of fibers used, the preform geometry, the type of process utilized fordensification, as well as the ultimate graphitization (heat-treat) temperature. The carbonfibers, which carry the majority of the mechanical and thermal load, are chosen on the basisof the desired final properties of the composite. Pitch-based carbon fibers are utilized for highmodulus up to (930 GPa) (Fibraplex) and high thermal conduc 1

applications while polyacrylonitrile (PAN)-based carbon fibers are used where highstrength (up to 6.96 GPa) (HexTow) is the desired property. Rayon and cellulose-basedcarbon fibers are required in ablative and insulation applications. The tow size (# offibers/tow) is specified on the basis of the desired properties and cost.

Fig: The Space Shuttle and its launch system require carbon carbon leading edges and anose-tip for the shuttle as well as carbon carbon nozzles and exit cones for the solid rocket

boosters

After the preform geometry has been selected and fabricated, it is then densified with acarbon matrix, which fills the space between the fibers and distributes the load among thefibers. For other types of composites the desired final matrix material is simply used to fill

the preform. However, in contrast, no carbon matrix material exists. It must be formed byconversion of some sort of hydrocarbon material to carbon by a process that involves a spe-cific type of pyrolysis known as carbonization. Thus, matrix precursor materials are placed inthe preform and then converted to carbon. This can involve a variety of processes, such as,chemical vapor deposition (CVD) process utilizing methane or propane at elevated tem-peratures , the charring of a polymer such as phenolic or polyimide as well as thepolymerization and coking of a pitch material to produce a high-quality graphitizable matrix.All of these processes have been studied extensively.

The first two processes are fairly well understood, but the third process has been verydifficult to elucidate until recently because of the large number of oligomeric species(~2000) in the pitch as well as the fact that the pitch material destroys analytical equipment.Carbon carbon composite aircraft brakes are manufactured by the CVD process. This processrequires a long period of time to densify a preform because it is necessary to keep thetemperature and gas concentration as low as possible in order to diffuse the hydrocarbon gasinto the preform before it deposits on the fiber surfaces. This process is limited to relativelythin (

Fig. Specific strength Vs temperature for composites and metals

After the hydrocarbon matrix is placed in the preform, it is carbonized to form the carbonmatrix. Depending on the type of precursor matrix material the carbon yield can vary from~50% (phenolic) to ~85% for a synthetic pitch. Obviously, even if the hydrocarbon matrixcompletely fills the preform (which it does not), several densification cycles are required.Depending on the application, either after each densification/carbonization cycle or after thelast cycle, the densified preform can be graphitized at temperatures in excess of 2200C inorder to improve the properties.

Although carbon carbon composites are excellent high-temperature structural materials, theirAchilles heel is oxidation. Above 425C, carbon starts to oxidize if unprotected limiting theuse of these composites for long-term high-temperature applications. Various oxidationprotection coating and inhibitor technologies have been tried, with most not being successful.A silicon carbide coating has performed well on the leading edges of the Space Shuttle wherethe temperature is below 1500C and a ZrB2-SiC coating performs well up to 1800C. Abovethis temperature only rhenium coated with iridium has proved to be a successful long-termcoating up to 2400C. Another way to decrease the oxidation rate is to densify the carbonfiber preform with a ceramic matrix, such as, silicon carbide producing a ceramic matrixcomposite.

Ceramic-Matrix Composite:

Since monolithic ceramics are brittle materials, even microscopic flaws can greatly reducethe strength of the component. This makes them unsuitable for many applications. However,the inherent brittleness of ceramic materials can be overcome by the use of continuous ordiscontinuous ceramic or carbon fiber reinforcements resulting in a ceramic matrix composite. This dispersed phase is designed to improve toughness by bridging the cracks and keepingthe material intact when it fractures. The fibers can also de-bond and slide through the matrix,dissipating energy and preventing fracture.

The reinforcement consists principally of fibers of alumina, silicon carbide, carbon, titaniumboride (TiB2), aluminum nitride (AlN), zirconium oxide (ZrO2), yttrium aluminum garnet(YAG), and alumina-silica (mullite), with SiC and carbon fibers being the most popular dueto their high strength (3 GPa - SiC) and modulus (270 GPa - SiC) (COI Ceramics). Thematrix by definition is a ceramic with the most important commercial matrices being SiC andAl2O3. To enhance oxidation resistance oxide (fiber)-oxide (matrix) composites are utilized.Depending on the constituents, these composites are able to function well in an oxidizingenvironment for long periods at temperatures up to 1400CThe ceramic fiber structures can be densified with the matrix material by using variousprocesses such as slurry infiltration, polymer impregnation and pyrolysis (PIP), chemicalvapor infiltration (CVI), melt infiltration (MI), electrophoretic deposition of a ceramic pow-der, as well as sol gel processing.

These composites are utilized in high-temperature oxidizing environments, such as inturbines, jet engines (flaps, vanes, seals, flame holders), heat shields, burner tubes, and heatexchangers .

Metal-Matrix Composites:

Although metal matrix composites (MMCs) do not function at as high a temperature ascarbon and ceramic-based composites, they function well where ductility as well as enhancedmechanical properties are required. These composites consist of a metal matrix (Al, Mg, Ti)reinforced with a continuous or discontinuous fiber (boron or carbon) or particulate (SiC, C,Al2O3, B4C). Because the matrix is metallic these composites also have good thermalconductivity. Since the vast majority of applications utilize an aluminum matrix, thesecomposites are usually employed in temperatures below 500C with most uses takingadvantage of the stiffness and low CTE of these composites. The principal applications above200C have been in the automotive industry for use as cylinder liners , connecting rods , as

and fatigue resistance, hardness, wear and abrasion resistance, as well as thermal shockresistance are required These composites are usually produced employing powder metallurgytechniques.

Modes of Failure at High Temperature

Some common modes of failure of materials are discussed below:

Melting and Softening

Mgenerally chemically very reactive with its environment. Thus, one needs high meltingcompounds for high-temperature applications. Impurities usually lower the melting point, anddeep eutectics can be disastrous. Glasses do not melt but they soften on going through theirglass transition. Glasses can crystallize, sometimes transforming a dense solid material to aporous powdery one. Even crystalline solids become soft and deformable at high temperatureand brittle materials become ductile . Such effects occur typically above about two-thirds ofthe melting point in degrees Kelvin.

Vaporization

Many solids vaporize (sublime) rather than melt. Others melt and then boil. At reducedpressure vaporization is more extensive. The vaporization behavior of metals and their oxidescan be quite different. For example, metals such as tungsten, molybdenum, rhenium, andosmium are refractory and useable to temperatures above 2000C, while their oxides are lowmelting and volatile. Their ready oxidation and the poor high-temperature behavior of theiroxides limit the use of such metals to non-oxidizing environments (inert gas or vacuum).

Corrosion and Chemical Reaction with the Atmosphere

Reactions include oxidation, nitridation, reaction with H2O and CO2 to form hydrates andcarbonates, and reaction with traces of HCl, SO2, and other corrosive gases. Oxidation,especially of small particles and dusts, can be explosive and reaction of active metals such aszirconium with water to form the metal oxide plus hydrogen can lead to hydrogen explosionsif air is also present. One should also note that nitrogen is not always a suitably inertgas for metals that readily form stable nitrides at high temperature, for example, titanium,zirconium, and tantalum.

Diffusion and Solid-State Reaction

When two different materials are in contact, they will diffuse into each other or form a layerof reacted material if such reaction is thermodynamically favorable. Grains will grow and thematerial may become porous and change in other physical properties. In composite structures,such as thermal barrier coatings, different layers may delaminate or material may spall off.On the other hand, adherent protective coatings of reacted (oxidized) metal or intentionallyadded other protective layers can greatly extend the use and lifetime of materials at hightemperature, for example, in protective oxide layers and thermal barrier coatings

Solid Solid Phase Transformations

If there are phase transitions among different crystal structures, the properties of the high-temperature phase may be quite different from those of the low-temperature phase. If thephase transition involves a large volume change, cracking or disintegration may occur.

MATERIALS:

Carbon-Carbon Composite:

Reinforced carbon-carbon composite has low density, a high strength to weight ratio, and theability to withstand high temperatures . Though Nimonic 263 is a high strength materialcompared to the carbon-carbon composite, its strength to density ratio is significantly lower.

Carbon-carbon composite retains its properties up to 2273 K without considerable decay instrength. Less brittle than many other ceramics, carbon-carbon composite have low impactresistances when compared to most non-ceramic materials. For this reason, carbon-carboncomposite is not used for the entire nose section. Carbon-carbon is a very costly materialexceeding $100,000 per panel so it is only used as a small strip at the leading edge of the nosecone. For reference, this material has been used for making the leading edge of the nose coneof NASA X43 hypersonic vehicle and Space Shuttles .

Tungsten:

Tungsten is used for the rest of the nose cone because of its strength, impact resistance,erosion resistance, corrosion resistance, resistance to thermal shock. Though titanium (Ti-6A1-4V) has high strength to weight ratio compared to tungsten, it was not used because itsstrength at high temperatures deteriorates much more quickly than Tungsten. Forreference,Tungsten was also used for the nose of NASA X-43 hypersonic vehicles.

SUMMARY AND CONCLUSION:

The major objective is to minimize the temperature generated by atmospheric friction and toreduce the drag on the body. Several materials can be used but since all these aerospace anddefence missions involve life and extremely high cost only the best possible materials shouldbe used

Considering the above mentioned factors it is inferred that Carbon-Carbon composite will beused as a small strip at the leading edge of the nose cone as it is very costly and for the restof the surface Tungsten can be used which has reasonably good strength at the highertemperatures.

Future Material:

Recent study has shown that Ultra High Temperature Ceramics can be potential replacementfor Carbon-Carbon composites providing cost advantage.

References:

1. Introduction to aerospace materials by Adrian P. Mouritz2. High Temperature Materials and Mechanisms Edited by Yoseph Bar Cohen3. Essentials of Materials Science and Engineering by Donald R. Askeland4. Project SENTINEL: Design of a Long-Range, High-Speed, Precision-Strike

Tactical Weapon by Vivek Ahuja, Jason Cary5. Ground Testing of Aerospace Vehicles Including Engines by A.K. Verma, K.N

Narayanan and George Joseph