IRIDIUM-COATED RHENIUM RADIATION-COOLED ROCKETS

Brian D. Reed, James A. Biaglow, and Steven J. Schneider

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135

SUMMARY

Radiation-cooled rockets are used for a range of low-thrust propulsion functions, including apogee insertion,

attitude control, and repositioning of satellites, reaction control of launch vehicles, and primary propulsion for

planetary spacecraft. The key to high performance and long lifetimes for radiation-cooled rockets is the chamber

temperature capability. The material system that is currently used for radiation-cooled rockets, a niobium alloy

(C103) with a fused silica coating, has a maximum operating temperature of 1370 °C. Temperature limitations of

C103 rockets force the use of fuel film cooling, which degrades rocket performance and, in some cases, imposes a

plume contamination issue from unburned fuel. A material system composed of a rhenium (Re) substrate and an

iridium (Ir) coating has demonstrated operation at high temperatures (2200 °C) and for long lifetimes (hours). The

added thermal margin afforded by iridium-coated rhenium (Ir/Re) allows reduction or elimination of fuel film

cooling. This, in turn, leads to higher performance and cleaner spacecraft environments. There are ongoing

government- and industry-sponsored efforts to develop flight Ir/Re engines, with the primary focus on 440-N,

apogee insertion engines. Complementing these Ir/Re engine development efforts is a program to address specific

concerns and fundamental characterization of the Ir/Re material system, including (1) development of Ir/Re rocket

fabrication methods, (2) establishment of critical Re mechanical property data, (3) development of reliable joining

methods, and (4) characterization of Ir/Re life-limiting mechanisms.

INTRODUCTION

Radiation-cooled rockets are used for a variety of low-thrust propulsion functions, including apogee insertion,

attitude control, and repositioning of satellites, reaction control for launch vehicles, and primary propulsion for

planetary spacecraft. The majority of radiation-cooled rockets use nitrogen tetroxide (NTO)/monomethylhydrazine

(MMH) or NTO/hydrazine (N2I-I4) bipropellant combinations. The key to high performance and long lifetimes for

radiation-cooled rockets is the thermal and oxidation limits of the combustion chamber.

The state-of-the-art (SOA) material system for radiation-cooled rockets is a niobium alloy (C 103) with a fused

silica coating (R512E), which has been in use for over 20 years. Although C103 has a high melting point, it is very

susceptible to oxidation, necessitating the protective silica coating. The silica coating, however, has a temperature

limitation of 1370 °C. Fuel film cooling is employed in C103 rockets to maintain this thermal limit. The use of

extensive amounts of fuel film cooling in low-thrust rocket engines significantly degrades performance and, in some

missions, represents a source of plume contamination.

A material system composed of a rhenium (Re) substrate and an iridium (Ir) oxidation-resistant coating offers

substantially higher temperature capabilities. Ir-coated Re (Ir/Re) rockets have a maximum operating temperature of

2200 °C, which for NTO/MMH or NTO/N2H4 propellants, allows for nearly optimal performance and extended

chamber lifetimes. Long-life operation (hours) of Ir/Re rockets, at temperatures from 1925 to 2200 °C, has been

demonstrated at 22-, 62-, and 440-N thrust levels (ref. 1), at performance levels significantly above SOA C103

rockets.

Despite demonstrations of Ir/Re rocket performance and life, the technology has not yet been used in flight

systems because of fabrication and engineering concerns. There are ongoing NASA- and industry-sponsored efforts

to develop flight Ir/Re engines, primarily for applications requiring 440-N thrust levels. Complementing these flight

engine development efforts is a program to address concerns that have arisen during development and to provide

fundamental characterization of the Ir/Re material system. This report will summarize the ongoing Ir/Re NASA

technology program.

NASA TM-107453 1

https://ntrs.nasa.gov/search.jsp?R=19970036365 2018-05-11T16:32:47+00:00Z
IX/ReDEVELOPMENTHISTORY

Inthemid-1980's, an effort was initiated to find a material system capable of sustained operation at 2200 °C

in oxidizing rocket combustion environments (ref. 2). Refractory metals, ceramics, composites, and carbon-carbon

materials were evaluated as substrate materials. Platinum (Pt) group metals, Engle-Brewer compounds, and ceramics

were considered as oxidation-resistant coatings. Monolithic Ix-Re-rhodium (Rh) alloys and ceramic/metal

(CERMET) alloys were also considered.

As a result of this material screening, a combination of Re (substrate) and Ir (oxidation-resistant coating) was

selected. Re has a high melting point (3175 °C), excellent high temperature strength, and does not have a ductile-tobrittle transition point, which makes it unique among refractory metals. Ir is oxidation resistant and has a temperature capability of 2400 °C and a coefficient of thermal expansion close to that of Re. Fabrication of IX/Re parts had

been demonstrated in an earlier Air Force technology program (ref. 3), in which chemical vapor deposition (CVD)

was used to coat a mandrel with lx and then Re. This CVD process will be described in greater detail later.

Under a NASA technology program with Gencorp Aerojet (ref. 2), 22-N demonstrator rockets were fabricated

(by using CVD) and tested on NTO/MMH propellants. One rocket was operated over 15 hr and 2700 cycles, at

2200 °C wall temperature, without failure. A flight-type, 22-N IX/Re engine demonstrated a specific impulse Isp of

310 sec, at an area ratio of 150:1. This performance is 20 sec above that provided by SOA 22-N, C103 engines. The

Ix/Re engine was also subjected to over 100 000 pulses without failure. Figure 1 shows the 22-N Ix/Re rocket firing

at a 2200 °C outer wall temperature.

In a follow-on NASA technology program (ref. 4), Aerojet demonstrated an Isp of 305 sec for a 62-N engine

(75:1 area ratio) on NTO/MMH. This again represented a 20-see increase in performance over a 62-N, C103 engine

of the same design. In the same technology program (ref. 5), a 440-N engine (286:1 area ratio) was tested for over

6 hr. This flight-type engine achieved 322 sec on NTO/MMH, a 12-see increase over 440-N class, C103 engines.

The production of 62- and 440-N engines also demonstrated the scalability of the process.

Ix/Re technology was proposed for the primary propulsion engine on the Mariner Mark II spacecraft (ref. 6). A

mass savings of 600 kg for the Comet Rendezvous and Asteroid Flyby (CRAF) mission was estimated through the

use of IX/Re technology. However, the state of IX/Re technology was considered immature at the time. As a result,

SOA C103 rockets were baselined on Mariner Mark II and ultimately used in the Cassini mission. IX/Re rockets

were proposed as drop-in replacements of the Shuttle Orbiter reaction control system thrusters (ref. 7). In this

proposed application increased lifetimes for the chambers, rather than performance, were the drivers.

Currently Ix/Re development efforts center on the production of high-performance, 440-N class rockets for the

next generation of apogee insertion engines for geostationary satellites. Ix/Re has the thermal margin to offer nearly

the maximum performance possible for NTO/MMH and NTO/N2H4 propellants. The same engine class also has

application to NASA planetary exploration missions. In the course of these efforts to develop flight IX/Re engines,

concerns have arisen about (1) the reliability and cost of It/Re rocket fabrication methods, (2) the integrity of the Re

mechanical property data base, (3) the ability to join Re chambers to dissimilar metals, and (4) the nature of IX/Re

life-limiting mechanisms. A NASA program was undertaken to address these concerns (summarized in fig. 2) in

order to aid the efforts to bring IX/Re technology to flight engines.

IX/Re FABRICATION PROCESSES

Because Re cannot be machined by using conventional metal working techniques, an IX/Re fabrication process

that is cost effective, reproducible, and lends itself easily to production-type operation is critical to the acceptance of

this technology for flight systems. Several IX/Re fabrication options are being explored as discussed in the following

sections.

Chemical Vapor Deposition

CVD is currently the only established process for fabricating IX/Re chambers. CVD is a method of plating which

relies on the chemical reaction of a vapor at a surface to form solid structural deposits (ref. 8). Because deposition

occurs on an atom-by-atom basis, purity levels above 99 percent are achievable. In CVD processes, gaseous precursors of IX and then Re are flowed over a heated mandrel that conforms to the contour of the chamber. Upon contact with the mandrel, thermal decomposition of the gaseous precursor occurs, which results in the deposition of the

material onto the mandrel. After the Ir and Re layers are deposited in this manner, the mandrel is then chemically

NASA TM-107453 2
removed, leavingafreestandingIr/Rechamber. Successful productionofCVDstructuresdependsonthedeposition

parameters, includingthegaseousprecursorfthematerial tobedeposited, themandrel material andtemperature,

thegasconcentration andflowrate,reactionchamber geometryandpressure, anddepositionthickness.

CVDhasbeenusedtoproduce22-to440-Nclasschambers, demonstrating thescalabilityoftheprocess. The

CVDprocessproducesdense,adherent, andhighpurityIr coatings. However, thereareconcernsabouttheintegrity

oftheCVDResubstrate. TheCVDResubstrateisdepositedinmultiplelayers(usuallythreeorfour).Whenheated

theindividualRelayerstendtoseparate, allowingslippage(ref.9).Thisphenomenon hasbeenshowntoleadto

somevariabilityofCVDRemechanical properties, whichishighlyundesirableforthedesignandincorporationof

flightengines.

PowderMetallurgyRe/Electroformed DepositedIr

Powdermetallurgy(PM)isanalternativefabricationprocessthatshowsignificant promise. InthePMprocess,

high-purityRepowderispressedandsintered, producingReingotsofabout95percentdensity.Thedensityofthe

pressedandsinteredReisthenincreasedtoabove99percentviaahotisostaticpressure(HIP)treatment. Afterthe

HIPtreatment, theReingotiselectrical dischargemachinedintothedesiredchamber contour.Anongoingtechnologyprogram(Privatecommunication fmM.J.HamisterregardingNASAContractNAS3-27738-I, July1996)

wouldmodifythePMprocesstoproducenear-net shapeRechambers, savingprocessingtimeandmaterial.It has

beenshownthatthePMReprocessprovidesrepeatablequalityandPMsampleshavetensilestrengthscomparable

toorabovethoseofCVDRe.ThePMprocessisonlyapplicabletotheResubstratefabrication. Aseparateprocess

isrequiredforapplicationoftheIr coating.

Electroformed depositionfED)hasbeenusedtoapplyIr coatingstoPMResubstrates(ref.10).IntheED

process, aplatingbathactsasanelectrolyticcell.Theworkpiecetobecoatedservesasthecathode, whiletheanode

iseitheranaqueousormoltensaltsolutionofthematerial tobeplated.Theworkpieceisimmersedinthebathand

currentispassedthroughtheanode,transferringionsofthematerial totheworkpiece. Todate,moltensaltplating

bathshavebeentheonlysuccessful methodofdepositingthickIr coatings. The22-and220-NclassPMRe

substrateshavebeencoatedbyusingEDandtestedonvariouspropellant combinations (refs.11and12).It hasbeen

foundthat,toproduceanadherent EDIr coating,it isnecessary toperformapost-densification HIPprocess

(ref.12).

MaturingIr/ReFabricationProcesses

Thereareothercandidateprocesses forfabricatingtheIr coatingorliner,theResubstrate, ortheentireIr/Re

chamber. Thesealternativeprocesses arecurrentlylessmaturethantheCVDandPMRe/EDIr.ChamberscomposedofathinCVDIr/Reshell,overwrappedwithacarbon-carbon fiberweave(whichbearsthemechanical loads)

havealsobeendemonstrated (ref.13).Thekeytotheacceptance ofanytheseIr/Refabricationalternativesistheir

abilitytoprovidereproducible, production-type qualitypartsatcostscomparabletoorlowerthanCVD and PM

Re/ED Ir.

Re MECHANICAL PROPERTY DATA BASE

Accurate data on Re mechanical properties are critical to the design and integration of It/Re engines. Tensile

and fatigue data are necessary for assessing the engine's ability to survive the stresses induced by the launch vibration environment. Data on annealed samples are particularly important, as they simulate the condition of the engine

after hot fire acceptance tests on the ground. Creep data are important in determining the engine's response to

multiple, long-duration on-orbit firings.

Currently the data base for Re mechanical properties is scattered. Reported data often do not include the fabrication or material history. This information is crucial to interpreting and applying the data. In particular, there is a

lack of data on the manufactured Re forms of interest for rocket applications. To address this data shortfall, a NASA

program was initiated to establish a Re mechanical database, with emphasis on CVD and HIP PM Re samples.

Tensile, low cycle fatigue, and creep data have been accumulated on Re both at room and elevated temperatures

(refs. 14, 15, and 16). Some of the accumulated mechanical data are summarized in tables I to III.

The data collected to date indicate that the properties of Re are strongly dependent on its fabrication method and

history. There was some variability in the CVD Re data, probably because of slippage between Re layers at elevated

NASA TM-107453 3
temperatures. HIP PM Re was characterized by sharp failures with almost no necking of the sample. CVD and HIP

PM Re samples completed 100 cycles of fatigue testing and demonstrated sufficient tensile strength for the rocket

application (ref. 15). More data, particularly data pertaining to low cycle fatigue, will be required for the proper

design of Re engines.

JOINING OF Re TO DISSIMILAR METALS

For the flight engines, the Re chambers need to be joined to injectors and nozzle skirts composed of dissimilar

metals. Common injector materials include 304L stainless steel (SS), titanium (Ti), and C103. Since Re is a heavy

material and high-temperature capability is not needed downstream of the thi'oat region, Re engines will employ

C103 nozzle skirts. Both the injector-to-chamber and the chamber-to-nozzle joints are subjected to thermal cycling

from 25 °C to a maximum of 1370 °C, with the number and duration of the thermal cycles dependent on the

application.

Because of its reliability and relative ease, electron beam (EB) welding would likely be the preferred method of

joining Re chambers to C103 nozzle skirts. However, attempts to EB weld Re directly to C103 have resulted in

brittle joints because of the formation of a Re-niobium (Nb) sigma phase (refs. 2 and 9). This brittle intermetallic

phase, which forms above 2162 °C, precludes the use of fusion methods in directly joining Re to C103. Brazing and

transition joints are alternatives to directly EB welding Re chambers to C103 skirts and will be discussed in the next

section.

Brazing

Brazing has been investigated extensively as an alternative to EB welding. Braze materials for joining Re to

304L SS, Hastelloy B2, and unalloyed Nb were evaluated by Aerojet under a NASA program (ref. 2). Palcusil 25,

Silcoro 75, Paisil 10, Nioro (82 gold (Au)-I 8 nickel (Ni)), and 50Au-50 copper (Cu) were evaluated for their wetting

ability. All of the braze alloys were found to have good to excellent wetting characteristics, with the exception of

Silcoro 75 and 50Au-50Cu for Nb. In lap and ring shear testing, Nioro joints were found to be the strongest. Re-to304L SS and Re-to-Hastelloy B2 joints were achieved by using a "parent metal braze" technique. Here only the

304L SS and Hastelloy B2 were melted, wetting the Re and solidifying to form the joint. This technique was used to

join a Re chamber to a 304L SS injector in a flight-type, 440-N engine (ref. 5).

TRW investigated methods of joining Re-to-C103 in a NASA program to develop a liquid oxygen/N2H4, 440-N

class engine (ref. 9). Several braze materials were evaluated for furnace brazing of Re-to-C 103, including 65Au-35

palladium (Pd), 50Au-50Pd, and Ti. These three braze joint samples were subjected to thermal aging and thermal

cycling tests. The Ti sample showed a small crack at the braze/Re interface, caused by the formation of a brittle

phase. The 50Au-50Pd sample showed void areas where there was poor flow of braze and shrinkage which formed

cracks. The 65Au-35Pd sample showed only some small voids and no diffusion of the braze material into Re or

C103. Diffusion of Nb and Hf (from the C103 alloy) into the braze was found, however. Based on this investigation,

the 65Au-35Pd braze alloy was recommended for furnace brazing Re-to-C103.

Although suitable braze materials have been found for joining Re to dissimilar metals, brazing has disadvantages for joining chambers to nozzle skirts. The brazing process often requires a complex design at the interfaces to

assure proper wetting of the joining area. Braze layers also have limited strength, temperature, and thermal cycling

capabilities. This calls into question the ability of braze layers to survive the rigors of launch vibration and multiple

on-orbit firings. An all-welded engine would eliminate these potential problems.

Transition Joints

An alternative to brazing is the use of a transition joint composed of Re and C103. The transition ring can be

formed by nonfusion methods. This avoids the formation of a brittle Re-Nb intermetallic, because bond line temperatures never reach 2162 °C. The Re chamber can then be EB welded to the Re side of the joint and the C103 skirt

can be EB welded to the C103 side. A transition joint would also be used for the injector-to-chamber joint. Nonfusion joining methods of forming transition pieces include diffusion bonding, explosive bonding, and inertia welding. In diffusion bonding, heat and pressure are applied uniformly to expand one piece into the other. Explosive

bonding is a cold welding process that uses controlled explosive detonations to join two pieces. Inertia welding uses

the stored energy in a flywheel to impart one piece into a stationary piece.

NASA TM-107453 4
FlatplatecouponsofRe-to-C103joints,formedbyexplosiveanddiffusionbondingwereevaluatedinaseries

ofsheartests(refs.16and17).Couponsweresubjectedtothermalcycling,thermal aging,andEBweldingof Re

and/orC103tabs.Ultimateshearstrengthwasusedasafigureofmeritforcomparisonfas-receivedcoupons. For

bothexplosive- anddiffusion-bonded coupons, thejointstrengthonlydegradedafterbeingsubjectedtothe

combinationofC103EBweldtabsandthermal cycling(100cycles,from100to1100°C).Thiscombination

apparentlyresultedinenoughinterdiffusionbetweenReandNbtoformanalloythatweakenedthejoint.This

experienceindicatesthatcaremustbetakenwhenEBweldingthechamber andnozzleskirttothetransitionjoint.

AnadequateheatsinkshouldbeprovidedduringEBweldingandthejointdesignshouldhaveageometrythat

wouldkeepthebondlinetemperaturebelow2162°C.

TRWevaluatedinertia-,diffusion-, andexplosive-bonded jointcouponsunderaNASAtechnologyprogram

(Privatecommunication fromM.L.ChazenregardingNASAContractNAS3-26246). InthistestingRewasnot

directlybondedtoCl03,sincetherewasconcernthatdiffusionofhafnium(Hf)fromCl03intoRewouldresultina

brittleHf-Reintermetallicphase.CouponsofRewerejoinedtoaNb-tantalum(Ta)alloybyinertiawelding,topure

Tabydiffusionbonding, andtoC103withaTainterlayerbyexplosivebonding.Afull-sizeringconfigurationwas

used;it providedamorerealisticassessment ofthejointintegritythanflatplatecoupons. Couponsweresheartested

afterbeingsubjectedtoshearloadingandthermal cyclingthatsimulatedthejoint'santicipateddutycycle.All ofthe

couponswerefoundtohavemorethansufficientmargintowithstandtheanticipateddutycycle.Basedonthisfullsizeringtesting,inertia-,diffusion-,andexplosive-bonding areallacceptablemethodsforformingtransitionjoints.

Ir/ReLIMITINGMECHANISMS

Ir/ReLifeModel

Life-limitingmechanisms forIr/RewereinvestigatedintheinitialIr/Rerocket technology programs (ref. 2).

The primary mechanism was found to be grain boundary diffusion of Re into the Ir layer. As the Re concentration at

the inner wall increases, the oxidation resistance of the Ir coating slowly degrades. At a critical Re concentration

level there is a significant increase in the oxidation rate, which results in loss of the Ir coating. By using this failure

mechanism, a first-order life model for Ir/Re rockets was established from Ir-Re diffusion and oxidation data. The

model is useful in assessing the ability of Ir/Re to survive in different combustion environments.

Thermogravimetric analyses (TGA) were performed at 1540 °C with Ir, Ir-20 percent Re, and Ir-40 percent

Re samples at 0.014-kPa oxygen partial pressure and with Ir and Ir-20 percent Re samples in 1.93-kPa oxygen

partial pressure (ref. 2). The TGA results and calculated material recession rates are shown in table IV. At

0.014-kPa oxygen pressure, the Ir oxidation rate increased significantly between Ir-20 percent Re (0.361 lam/hr) and

It-40 percent Re (58.1 I.tm/hr). Based on these data, the failure criterion for Ir/Re rockets was established empirically

as 20 percent Re at the Ir inner surface.

The phase diagram for Ir-Re is shown in figure 3. Here the lower abscissa is weight percentage Re and the

upper abscissa is atomic percentage Re. The dashed lines in the phase diagram represent uncertainty associated with

the locations of the region boundaries. The It-Re phase diagram shows that the Ir-20 percent Re sample was a

single-phase solid solution, while the It-40 percent Re sample was a two-phase mixture. The two-phase region

boundary represents the solubility limit of Re in Ir. The Re-rich phase (which was reported to have a composition of

29 percent Ir and 71 percent Re) accounted for over a third of the Ir-40 percent Re alloy at 1540 °C. The sudden

large increase in oxidation rate between the two previous samples was attributed to the Re-rich phase, which was

thought to be more susceptible to oxidation (ref. 2). Based on these data, the failure criterion for It/Re rockets is

probably better set at the Re concentration corresponding to entry into this two-phase region.

Furnace oxidation testing of vacuum arc melted It-Re buttons was conducted to test the above hypothesis

(ref. 18). Samples were tested at 1500 °C in oxygen partial pressures of 0.69 and 6.90 kPa. Because of porosity in

many of the samples, there was a large amount of anomalous oxidation data. The remaining acceptable data (at

0.69-kPa oxygen partial pressure) from the study, that is, data from samples with little or no porosity, is shown in

table V. It can be seen that there is an order of magnitude increase in oxidation rate from 19.8 percent (0.369 Ixm/hr)

to 27.6 percent (4.29 Ixm/hr) Re concentration. The sparsity of acceptable data between 20 percent and 30 percent Re

concentrations, however, leaves uncertainty as to whether the oxidation rate increase can be attributed to entry into

the two-phase region of the It-Re phase diagram (fig. 3).

Diffusion data were collected from CVD It-Re samples that had been vacuum annealed at 1400, 1700, and

1900 °C for 8 hr each (ref. 19). Electron microprobe analyses of the samples revealed significant diffusion of Re

along Ir grain boundaries, while there was little diffusion of Ir into the Re. Diffusion constants for the samples were

NASA TM-107453 5
determinedfrom curve fits of the microprobe line scans, with the assumption that there was only diffusion of Re into

the Ir layer. As expected, the diffusion constant was found to have an Arrhenius dependence on temperature T and

could be fitted to an equation of the following form:

D = D Oexp (-E/kT)

where D O = 3.14×108 cm2/sec and E/k = 1.42×104 K. The activation energy for diffusion E was found to be

1.23 eV.

A first-order life prediction model was established for Ir/Re rockets (ref. 20) by using the diffusion rate data in

the one-dimensional diffusion equation

OC/0t = D(_)C2/Ox 2)

where C is the Re concentration, t is time, and x is the position normal to the Ir/Re interface. This equation was

solved numerically by using the Crank-Nicholson differencing scheme, with the boundary conditions of zero Re flux

at the Ir (inner) surface and 100 percent Re concentration at the Ir/Re interface. The inner surface was assumed to be

removed at a fixed recession rate based on the TGA data at 1540 °C and 0.014 kPa. The furnace data gathered at

1500 °C and 0.69 kPa could also be used as the basis of the recession rate in the model. The failure criterion was set

at 20 percent Re concentration at the inner wall.

This Ir/Re life model is only as valid as the diffusion and oxidation data that it is based on. The diffusion rate

data have a predictable dependence on temperature and can probably be extrapolated to higher temperatures with

confidence (ref. 20). Oxidation rate data, however, are a strong function of temperature and oxygen partial pressure

and cannot be confidently extrapolated. Therefore the life model is only valid within the range of oxygen pressures

and temperatures for which Ir-Re oxidation rate data exist. Life model predictions at higher temperatures and

oxygen pressures lead to unrealistically long chamber lifetimes. This limits the use of the life model to 1540 °C,

until Ir-Re oxidation rate data at higher temperatures can be obtained.

Figure 4 shows the predicted chamber lifetimes for Ir/Re rockets, with Ir coating thicknesses of 50.8 and

101.6 txm (typical of Ir/Re rockets) at 1540 °C. Also plotted in the figure are the chamber lifetimes for C103 rockets,

based on rocket engine qualification tests conducted by various rocket manufacturers. An Ir/Re rocket with a

50.8-ktm Ir coating was predicted to have an order of magnitude increase in chamber lifetime over that of C103

rockets. Typically Ir/Re rockets would be operated at temperatures higher than 1540 °C for increased performance.

If Ir/Re rockets were operated at 1540 °C, performance gains would be sacrificed in lieu of greatly extended chamber lifetimes as in the proposed Shuttle thruster application described earlier (ref. 7). For comparison, a 22-N Ir/Re

rocket tested on NTO/MMH propellants (ref. 2) is also shown in figure 4. The rocket had a 50.8-_tm Ir thickness and

did not experience any erosion or degradation during testing.

Ir Degradation

In the initial Ir/Re technology programs with Aerojet (22-, 62-, and 440-N engines on NTO/MMH), no significant erosion or degradation of the Ir coating was observed. In this early testing, Aerojet used either a cooled

section between the chamber and injector or a Pt-Rh liner installed in the head end region of the chamber (refs. 2, 4,

and 5). In more recent testing by other rocket companies, Ir/Re chambers have been directly attached to injectors. In

this configuration, degradation of the Ir layer in the form of pitting was often experienced in the head end (nearinjector) region of the chamber. A reported case of this type of Ir degradation occurred in testing 22-N engines with

both NTO/MMH (after 1500 see of testing) and NTO/N2H 4 (after 450 see of testing) (ref. 21).

In the near-injector region of the chamber, it was likely that the flow field was still mixing with oxidizers and

combustion radicals present at the wall. In the instances when Ir/Re chambers were run with a cooled-section or

liner, the Ir was not exposed to the flow field mixing and combustion processes were complete. The circumstantial

evidence, then, would seem to point to exposure of the Ir layer to a mixing, combusting flow field as the cause of the

Ir degradation. The cause of the Ir degradation could be due to Ir reactivity with a combustion radical, oxidation of

the underlying rhenium through the Ir grain boundaries, or high partial pressures of oxidizers at the wall directly

oxidizing the Ir.

The reactivity of Ir and Pt in a fuel-rich air/hydrogen combustion environment has been investigated (ref. 22).

Laser-induced fluorescence measurements of Ir and Pt samples found hydroxyl (OH) concentrations significantly

above equilibrium. The measurements suggested that atomic oxygen and atomic hydrogen concentrations near the

NASA TM-107453 6
material surfacewerealsoabovequilibrium. OHisanintermediateproductinNTO/MMHandNTO/N2H4combustionenvironments. However, nosignificant differenceintheconcentration profileswasfoundbetweentheIr and

Ptsamples. Thesuperequilibriumconcentrations, then,maynothavebeentheresultofthereactionwiththematerial

surface, butacharacteristic ofthecombustionenvironment. Beyondthiswork,therehasbeennotestingconducted

toconclusivelydeterminethecausesoftheIrdegradation.

AlthoughthedirectcausesoftheIr degradationarenotwellunderstood, thephenomenon canbeavoidedby

onlyexposingtheIrlayertoawell-mixed, well-combusted flowfield.ThiswouldinvolveprotectingtheIrlayerin

therelativelylow-temperature near-injector regionwithalow-reactivitylinerorcoating.Thedegradationhasnot

occurredinthethroatregion,wherethehigh-temperature capabilityofIrismostneeded.

CONCLUSIONS

Ir/Rechamber materialsarekeyto high-performance, long-life, radiation-cooled rocket engines. There has been

extensive testing of Ir/Re rockets, on different propellant combinations, at thrust levels up to 440 N. A low-cost,

reproducible fabrication process for Ir/Re is critical for its acceptance in flight systems. CVD is the established

fabrication process for Ir/Re, although PM Re/ED Ir represents a viable alternative. It has been found that Re

mechanical properties are highly dependent on its fabrication process and history. Direct EB welding of Re to C103

nozzle skirts is not possible, because a brittle Re-Nb intermetallic forms at 2162 °C. Alternative Re-to-Cl03 joining

methods include brazing and transition rings formed by nonfusion methods, such as explosive-, diffusion-, and

inertia-bonding. The life-limiting mechanism for Ir/Re rockets is the grain boundary diffusion of Re into the Ir layer

and the oxidation of the resulting Ir-Re alloy from the inner surface. This mechanism is relatively slow, providing

for long Ir/Re chamber lifetimes. Ir degradation experienced in the near-injector region of some Ir/Re rockets was

probably the result of exposing Ir to a mixing, combusting flow field. This problem can be avoided by protecting the

Ir layer in the (relatively low-temperature) near-injector region with a low-reactivity liner or coating.

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3. Harding, J.T.; Tuffias, R.H.; and Kaplan, R.B.: Platinum Group Coatings for Refractory Metals. AFRPL-TR84-035 (AD-AI43792), Air Force Rocket Propulsion Laboratory, 1984.

4. Jassowski, D.M.; and Gage, M.L.: Advanced Small Rocket Chambers. Option 1:14 lbf Ir-Re Rocket. Final

Report. NASA CR-191014, 1992.

5. Jassowski, D.M.; and Schoenman, L.: Advanced Small Rocket Chambers, Option 3: I l0 lbf Ir-Re Rocket. Final

Report. NASA CR-195435, vols. 1 and 2, 1995.

6. Appel, M.A., et al.: Feasibility Demonstration of a 445N High-Performance Rocket Engine. Presented at the

1989 JANNAF Propulsion Meeting, voi. 1, 1989, pp. 445-455.

7. Krohn, D.D.: Space Shuttle Vernier Thruster Long-Life Chamber Development. AIAA Paper 90-2744, 1990.

8. Harding, J.T.; Kazaroff, J.M.; and Appel, M.A.: Iridium-Coated Rhenium Thrusters by CVD. NASA

TM-101309, 1988.

9. Chazen, M.L; Mueller, T.; Rust, T.: Space Storable Rocket Technology (SSRT) Program. Option 1: Program.

Final Report. NASA CR-191171, 1993.

10. Malone, G.A.; and Walech, T.: High Temperature Barrier Coatings for Refractory Metals. NASA CR-195475,

1995.

11. Chazen, M.L.; and Sicher, D.: High Pressure Earth Storable Rocket Technology Program Summary. AIAA

Paper 95-2939, 1995.

12. Reed, B.D.; and Dickerson, R.: Testing of Electroformed Deposited Iridium/Powder Metallurgy Rhenium

Rockets. NASA TM-107172, 1995.

13. Tuffias, R.H.; Williams, B.E.; and Kaplan, R.B.: Lightweight, Inexpensive Radiation-Cooled Advanced

Composite Combustion Chambers. AIAA Paper 95-2400, 1995.

14. Chazen, M.L.: Materials Property Test Results of Rhenium. AIAA Paper 95-2938, 1995.

15. Biaglow, J.A.: Rhenium Material Properties. AIAA Paper 95-2398 (NASA TM-107043), 1995.

NASA TM-107453 7
16. Reed, B.D.; and Biaglow, J.A.: Rhenium Mechanical Properties and Joining Technology. AIAA Paper

96--2598 (NASA TM-107317), 1996.

17. Reed, B.D.; and Morren, S.H.: Evaluation of Rhenium Joining Methods. AIAA Paper 95-2397 (NASA

TM-107055), 1995.

18. Reed, B.D.: High-Temperature Oxidation Behavior of Iridium-Rhenium Alloys. AIAA Paper 94-2893 (NASA

TM-106720), 1994.

19. Hamilton, J.C., et al.: Diffusion Mechanisms in Ir-Coated Re for High-Temperature, Radiation-Cooled Rocket

Thrusters. AIAA Paper 91-2215, 1991.

20. Jassowski, D.M.: Advanced Small Rocket Chambers. Basic Program and Option 2: Fundamental Processes and

Material Evaluation. NASA CR-195349, 1993.

21. Wood, R.S.: Experiences With High Temperature Materials for Small Thrusters. AIAA Paper 93-1962, 1993.

22. Barlow, R.S., et al.: Gas-Phase Measurements of Combustion Interaction With Materials for Radiation-Cooled

Chambers. AIAA Paper 91-2216, 1991.

TABLE L---CHEMICAL VAPOR DEPOSITED RHENIUM TENSILE

DATA (REF. 15)

Annealing Annealing Test 0.2 percent Ultimate

temperature, time, temperature, yield strength, strength,

°C hr oC MPa MPa

None None 25 310 674

None None 25 297 663

1482 0.5 25 124 674

1482 0.5 815 359 430

1482 0.5 815 317 440

1482 0.5 1370 200 367

1482 0.5 1370 138 146

1482 0.5 1370 193 344

1650 1.0 25 307 688

1650 1.0 25 285 698

1650 1.0 815 400 465

1650 1.0 815 390 445

1650 1.0 1370 345 456

1650 1.0 1370 267 290

NASA TM-107453 8
TABLE II.--POWDER METALLURGY

RHENIUM TENSILE DATA

HOT ISOSTATIC PRESSURE)

'REFS. 15 AND 16)

Annealing Annealing Test 0.2 percent Ultimate

temperature, time, temperature, yield strength, strength,

°C hr °C MPa MPa

1650 0.5 25 237 911

1650 0.5 25 232 916

1650 0.5 815 254 562

1650 0.5 815 264 498

1650 0.5 1370 180 216

1650 0.5 1370 191 252

None None 1370 208 217

None None 1370 211 220

None None 1370 171 202

None None 1370 173 198

None None 1510 143 151

None None 1510 138 150

None None 1650 104 111

None None 1650 101 107

None None 1860 105 113

None None 1925 67.4 73.0

None None 1925 78.8 79.4

None None 1925 87.6 98.2

None None 1925 78.1 92.5

None None 2090 71.0 80.2
TABLEIII._RHENIUMLOWCYCLEFATIGUEDATA

(REF. 15)

Fabrication Test

method temperature,

oC

CVD 25

CVD 25

CVD 815

CVD 815

CVD 815

CVD 815

CVD 1370

CVD 1370

HIP PM 815

HIP PM 815

HIP PM 1370

HIP PM 1370

aFailed on first cycle.

Stress,

MPa

152

152

276

276

276

276

138

138

138

128

131

138

Tension/ Cycles

compression completed

152 100

152 100

179 (a)

138 100

138 11

138 12

69 85

69 37

69 103

62 It20

62 100

69 100

bSample was accidenfly loaded to 248 MPa in pretest run.

TABLE IV.--THERMOGRAVIMETRIC

ANALYSIS OF IRIDIUM-RHENIUM

ALLOYS AT 1540 °(2

Rhenium Oxygen

concentration partial

preSStLr¢,

kPa

0 (pure It) 0.014

20 0.014

40 0.014

0 (pure It) 1.93

20 1.93

REF. 2)

Recession

rate,

0.153

0.361

58.1

0.865

2.44

TABLE V.--FURNACE OXIDATION

TESTING OF IRIDIUM-RHENIUM

ALLOYS AT 1500 °C AND 0.69-kPa

OXYGEN PRESSURE (REF. 18)

Rhenium Recession rate,

concentration Ism/hr

1.14 0.042

3.08 0.084

3.61 0.059

16.9 0.236

19.8 0.369

27.6 4.29

32.0 7.57

34.8 20.8

10
Figure 1.--22-N iridium-coated rhenium rocket firing.

11
Fabrication Technology Rhenium Mechanical Properties

Low-cost, reproducible

fabrication processes

• Tensile

• Low cycle fatigue

• Creep

Joining Technology

• Injector to chamber

• Chamber to C103 skirt

Re diffusion Ir =

f (temperature) __\\

/ / \

/ / \

Re/Ir diffusion zone --/ / \

z_ Oxidation of Ir = "_ Erosion of pitting of Ir in

f (temperature, oxygen heed end region

partial pressure, Re -Exposure to mixing,

concentration), combusting flow

• Re diffusion

• Ir/Re oxidation

• Ir reactivity

Figure 2.oNASA Research & Technology Program for iridium-coated rhenium chambers.

12
Recontent,at%

0 10 20 30 40 50 60 70 80 90 100

3500 I I I I I r I I I I

3000 --

2500 _-"_

_ 2447

2000 --

a)

Q.

E

0

1500 --

1000 --

500

0

Ir

_ _--s 3186 °C

2800 _+25 °C _-

.quid _ - -

.....,=. _ _J _1'-39 55 _|

I I

/ I

I I

(Ir) I I (Re)

I L

I I

I |

I l

I l

I It I I ] i I I I ]

10 20 30 40 50 60 70 80 90 100

Re

Re content, wt %

Figure 3.--Iridium-rhenium phase diagram.

p-

_=

o

O

.E

l1000

100

10

O Ir/Relayerlifethickness model prediction = 101.6 I_m

_-- O Ir/Re life model prediction

--------_'__----_" Ir layer thickness = 50.8 i_m

"_any A [] 22-N, Ir/Re rocket \\

-- / \ tested on MMH/NTO _

__ / \

_- Si/Nb, Company B

l_ Si/Nb, Company C

I I I I\ I I I I I I

1200 1300 1400 1500 1600 1700 1800 1900 2000 2100 2200

Temperature, °C

Figure 4.--Time to failure versus temperature for C103 and iridium-coated rhenium rockets.

.1

I

2300

13
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REPORT DOCUMENTATION PAGE OMB No 0704-0188

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1. AGENCY USE ONLY (Leave blank) 2. REPORT DATE 3. REPORT TYPE AND DATES COVERED

July 1997 Technical Memorandum

4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Iridium-Coated Rhenium Radiation-Cooled Rockets

6. AUTHOR(S)

Brian D. Reed, James A. Biaglow, and Steven J. Schneider

7. PERFORMINGORGANIZATIONNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Lewis Research Center

Cleveland, Ohio 44135-3191

9. SPONSORING/MONITORINGAGENCYNAME(S)ANDADDRESS(ES)

National Aeronautics and Space Administration

Washington, DC 20546- 0001

WU-632-1B-IB

8. PERFORMING ORGANIZATION

REPORT NUMBER

E-10736

10. SPONSORING/MONffORING

AGENCY REPORT NUMBER

NASA TM- 107453

11. SUPPLEMENTARY NOTES

Prepared forthelnternationalSymposium on Rhenium and Rhenium AlloyssponsoredbytheMinerals, Memls, and

Materials Society, Orlando, Florida, February9-13,1997. Responsible pe_on, BrianD. Reed, organization code5430,

(216)977-7489.

12a. DISTRIBUTION/AVAILABILITY STATEMENT

Unclassified - Unlimited

Subject Categories 20 and 26

This publication is available from the NASA Center for AeroSpace Information, (301) 621-0390.

12b. DISTRIBUTION CODE

13. ABSTRACT (Maximum200 words)

Radiation-cooled rockets are used for a range of low-thrust propulsion functions, including apogee insertion, attitude

control, and repositioning of satellites, reaction control of launch vehicles, and primary propulsion for planetary spacecraft. The key to high performance and long lifetimes for radiation-cooled rockets is the chamber temperature capability.

The material system that is currently used for radiation-cooled rockets, a niobium alloy (C103) with a fused silica

coating, has a maximum operating temperature of 1370 °C. Temperature limitations of C 103 rockets force the use of fuel

film cooling, which degrades rocket performance and, in some cases, imposes a plume contamination issue from

unburned fuel. A material system composed of a rhenium (Re) substrate and an iridium Or) coating has demonstrated

operation at high temperatures (2200 °C) and for long lifetimes (hours). The added thermal margin afforded by iridiumcoated rhenium (k/Re) allows reduction or elimination of fuel film cooling. This, in turn, leads to higher performance

and cleaner spacecraft environments. There are ongoing government- and industry-sponsored efforts to develop flight Ir/

Re engines, with the primary focus on 440-N, apogee insertion engines. Complementing these Ir/Re engine development

efforts is a program to address specific concerns and fundamental characterization of the Ir/Re material system, including

(1) development of Ir/Re rocket fabrication methods, (2) establishment of critical Re mechanical property data, (3)

development of reliable joining methods, and (4) characterization of Ir/Re life-limiting mechanisms.

14. SUBJECT TERMS

Iridium; Rhenium, Radiation-cooled rockets; Rhenium joining; Iridium/rhenium

fabrication; rhenium properties; Satellite propulsion

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