NASA/TM-2010-216233

Thermal Expansion of Vacuum

Plasma Sprayed Coatings

S. V. Raj and A.R. Palczer

Glenn Research Center, Cleveland, Ohio

March 2010
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NASA/TM-2010-216233

Thermal Expansion of Vacuum

Plasma Sprayed Coatings

S. V. Raj and A.R. Palczer

Glenn Research Center, Cleveland, Ohio

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

March 2010
Acknowledgments

The authors thank Dr. David Ellis for his comments and suggestions, and Mr. George Thom, Plasma Processes, Inc.,

Huntsville, Alabama, for vacuum plasma spraying the specimens.

Trade names and trademarks are used in this report for identification

only. Their usage does not constitute an official endorsement,

either expressed or implied, by the National Aeronautics and

Space Administration.

Level of Review: This material has been technically reviewed by technical management

Available from

NASA Center for Aerospace Information National Technical Information Service

7115 Standard Drive 5301 Shawnee Road

Hanover, MD 21076-1320 Alexandria, VA 22312

Available electronically at http://gltrs.grc.nasa.gov
Thermal Expansion of Vacuum Plasma Sprayed Coatings

S.V. Raj and A.R. Palczer

National Aeronautics and Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Abstract

Metallic Cu-8%Cr, Cu-26%Cr, Cu-8%Cr-1%Al, NiAI and NiCrAlY monolithic coatings were

fabricated by vacuum plasma spray deposition processes for thermal expansion property measurements

between 293 and 1223 K. The corrected thermal expansion, (AL/Lo)therma1, Varies with the absolute

temperature, T, as

(AL/L.o)the.a1= A(T – 293) 3 + B(T – 293)2 + C(T – 293) + D

where, A, B, C and D are regression constants. Excellent reproducibility was observed for all the coatings

except for data obtained on the Cu-8%Cr and Cu-26%Cr coatings in the first heat-up cycle, which

deviated from those determined in the subsequent cycles. This deviation is attributed to the presence of

residual stresses developed during the spraying of the coatings, which are relieved after the first heat-up

cycle. In the cases of Cu-8%Cr and NiAl, the thermal expansion data were observed to be reproducible

for three specimens. The linear expansion data for Cu-8%Cr and Cu-26%Cr agree extremely well with

rule of mixture (ROM) predictions. Comparison of the data for the Cu-8%Cr coating with literature data

for Cr and Cu revealed that the thermal expansion behavior of this alloy is determined by the Cu-rich

matrix. The data for NiAI and NiCrAlY are in excellent agreement with published results irrespective of

composition and the methods used for processing the materials. The implications of these results on

coating GRCop-84 copper alloy combustor liners for reusable launch vehicles are discussed.

1.0 Introduction

Many engineering components experience temperature variations during service, which often require

the availability of good quality data to ensure that they can be designed reliably and safely to meet the

required design objectives. The necessity for generating high quality data is more critical for aerospace

components especially those where two or more different materials are bonded to each other to fabricate

components with optimized design properties. For example, differences between the thermophysical and

mechanical properties of coating materials and the substrate can lead to the development of large residual

stresses, which can either distort or debond the coating from the substrate during thermal cycling of a

coated component during service.

The National Aeronautics and Space Administration (NASA) has been developing technologies for a

new generation of advanced reusable launch vehicles (RLVs) in its efforts to increase its future heavy lift

capacity in a more reliable and economical manner compared to the Space Shuttle. Combustion chamber

liners in liquid hydrogen (LHZ) fueled-rocket engines experience extreme conditions due to a combination

of environmental and thermo-mechanical effects. The combustion flame temperatures in the chamber

interior are about 3600 K whereas the backside of the approximately 1 mm thick liner wall experiences

cryogenic temperatures of 20 K (Refs. 1 to 6). Uncoated copper and its alloys have been used as

combustor liner materials in these regenerative rocket engines because their high thermal conductivities

enable efficient heat transfer from the combustion flame to preheat the cryogenic LH2 flowing in the

cooling channels. However, uncoated copper alloy liners undergo environmental degradation due to a

combination of the spallation of the locally formed copper oxide scales and "blanching," which consists

of repeated oxidation of the copper matrix and subsequent reduction of the oxide scale (Ref. 6).

NASA/TM-2010-216233
The design of the next generation of RLVs calls for using GRCop-84 (Cu-8(at.%)Cr-4%Nb) copper

alloy liners due to its superior properties compared to conventional high conductivity copper alloys, such

as NARloy-Z (Refs. 7 to 9). The application of protective coatings on GRCop-84 and other copper alloy

substrates can minimize or eliminate many of the problems experienced by uncoated liners and

significantly extend their operational lives in RLVs. This factor potentially translates to increased

component reliability, shorter depot maintenance time and lower operational cost. In addition, the use of a

suitable top coat to act as a thermal barrier' can allow the engine to run at hi gher temperatures thereby

resulting in increased thermal efficiency. As a result, several types of ceramic (Refs. 1 and 5) and metallic

(Refs. 10 to 15) coatings have been advocated as protective coatings for copper alloy liners. However, the

presence of a coating will significantly reduce the thermal conductivity of the liner compared to the

uncoated copper alloys, which suggests that metallic coatings are preferred to ceramic coatings.

As mentioned above, the application of coatings on a substrate can result in the development of

residual stresses whenever the coated component experiences variations in temperature. These

temperature variations can occur during different stages of processing, such as when the sprayed

component cools to ambient after vacuum plasma spraying as well as after hot isostatic pressing (HIP) of

the component. Differential thermal expansion can also result in thermally induced residual stresses

during temperature cycles in service. It is important to note that the surface temperature of the combustion

liner is expected to increase from cryogenic temperatures to 800 to 1100 K within 3 sec on starting the

engine. As a result, an enormous amount of differential thermal strain can develop in the coated substrate

and lead to coating spallation during the transient heat-up portion of the cycle. In order to predict the

performance of these coatings in service, as well as to gain insights on the nature and magnitudes of these

residual stresses, it is essential to conduct finite element analysis (FEA) to simulate the performance of

coated liners in a realistic combustion environment. The reliability of these analyses requires the

generation of high quality mechanical and thermophysical property data. Recently, it was demonstrated

that CuCrAI and N1CrAlY coatings deposited either by the cold spray or the vacuum plasma spray

techniques are potentially viable coatings for GRCop-84 combustion liners (Refs. 14 and 15). However,

thermophysical data for these sprayed coatings are either limited or nonexistent in the temperature range

of interest for use in RLVs. Although thermal conductivity data of vacuum plasma sprayed N1CrAlY

alloys (Refs. 16 and 17) have been previously reported in the literature, it is important to note that these

properties are sensitive to coinpositional and processin g variables. Thus, it is essential that

thermophysical data be generated on coatings sprayed under processing conditions and for compositions

similar to those developed for spraying the GRCop-84 liners in order to ensure reliable design models to

be developed.

The objective of this paper is to report the temperature dependence of thermal expansion for vacuum

plasma sprayed coatings Cu-Cr, CUCrAI, NiAI and NiCrAIY coating alloys between 293 and 1223 K. The

data are analyzed to obtain valuable insights on the behavior of these alloys during heat-up and cool-down

thermal cycles.

2.0 Experimental Procedures

2.1 Alloy Composition and Processing

Gas atomized copper alloy powders were procured from Crucible Research, Inc., Pittsburgh,

Pennsylvania, whereas the NiAI and NiCrAIY powders were obtained from Homogenous Metals, Inc.,

New York, and Sultzer Metco, Michigan, respectively. The nominal compositions of the alloy powders

were Cu-8(wt.%)Cr, Cu-26(wt.%)Cr, Cu-8(wt.%)Cr-1%Al, Ni-31.5(wt.%)Al (N1AI), Ni-17(wt.%)Cr6%A1-0.5%Y. Monolithic cylindrical coatings, typically 175 to 250 mm long and 19 to 25 mm thick,

were fabricated by spraying the powders on rotating steel mandrels, approximately 25 min in diameter, by

'The term "thermal barrier" is used to signify that the coated copper liners are likely to possess lower thermal

conductivity than uncoated copper alloys.

NASA/TM-2010-216233
the vacuum plasma spray (VPS) method. The Cu-8%Cr, Cu-8%Cr- 1 %Al, NiAI and NiCrAIY coatings

were deposited by the vacuum plasma spray method at Plasma Processes, Inc., Huntsville, Alabama. The

coated mandrels were hot isostatically pressed (HIP) between 1073 and 1273 K under argon gas pressures

varying between 100 and 210 MPa for times varying between 1 and 4 hr.

2.2 Thermal Expansion Measurements

Specimens were machined from the HIPed cylinders by electrodischarge machining (EDM). The

specimen dimensions for thermal expansion measurements were 25.4 mm long and 6.35 mm in diameter

for all coatings except for the Cu-8Cr-lAl alloy specimens, which were 12.7 mm long and 6.35 mm in

diameter. The thermal expansion of the specimens was measured between 293 and 1223 K using a

NETZSCH Dilatometer Model DIL 4020, which uses a single head design and high purity alumina as a

calibration standard. Measurements were made over three heat-cool cycles to minimize the effects of

compositional, microstructural and processing inhomogeneities on the data, verify their reproducibility,

and statistically average the regression coefficients. The specimen was placed in a sample holder and

aligned with a single push-rod with an applied constant load of 0.2 N. The specimens were heated from

293 to 1223 K at 10 K/min. in a He atmosphere flowing at 60 cc.-'min in the first cycle and furnace cooled

to 373 K in the first cool-down cycle. Subsequent cycles consisted of heating and cooling between 373

and 1223 K. The length changes were recorded by a computerized data acquisition system. The

experimental strain, DULp, where AL is the differential change in length, L-Lo, and Lo is the original

length of the specimen at room temperature, were measured. Specimen-to-specimen reproducibility

measurements were made on three different specimens in the cases of Cu-8%Cr and N1Al.

3.0 Results and Discussion

3.1 Microstructures

The microstructures of the monolithic coatings were nearly fully dense with little visible porosity

under an optical microscope. The Cu-26%Cr (Fig. 1(a)) coatings revealed second phase a.-Cr particles in

the matrix at higher magnifications (Fig. 1(b)). The NiCrAlY coating was also nearly fully dense although

fine grain boundary porosity was observed at the higher magnifications (> 400 X). The Cu-8%Cr coating

was nearly fully dense with very little porosity visible in the microstructures. Similar nearly full dense

microstructures were observed in the NiAI and Cu-8%Cr-1%Al coatings.

Figure 1.—(a) Low and (b) high magnification views of the optical microstructures of vacuum plasma sprayed Cu26%Cr monolithic coatings.

NASA/TM-2010-216233
3.2 Thermal Expansion

3.2.1 Cu-8%Cr

Figures 2(a) to (c) compare the variation of AL/Lo with T, during three heat-up and cool-down cycles

for three Cu-8%Cr specimens. For all three specimens, the magnitudes of thermal expansion in the first

heat-up cycle were higher than those recorded in subsequent heat-up and cool-down cycles. In contrast,

the thermal expansion was nominally identical from the first to the third cool-down cycle although the

data for specimen 3 exhibited slightly more scatter (Fig. 2(c)). This observed difference in the magnitudes

of AL/Lo between the first and subsequent cycles is most likely due to the presence of multiaxial residual

stresses initially developed in the splat-cooled grains of the coating microstructure during processing. The

nature and calculation of these residual stresses are fairly involved (Refs. 18 to 20). As a result, the

measured values of (AL/Lo)fi,cycle for the first heat-up cycle consist of a uniaxial residual strain

component and a thermal strain component. Thus,

(AL/LO)fi,5t cycle = (ALIT-o)themal + (AL/Lo)residual (1)

where (AL/Lo)tl,er,nal and (AL/Lo),.id„al are the thermal and residual strains, respectively. Since the

temperature dependence of AL/Lo for the first to the third cool-down cycles are nearly identical (Fig. 2(a)

to (c)), it is reasonable to assume that the magnitude of (AL/Lo)tl,en„al can be determined from these data.

Subsequently, the temperature dependence of (AL/Lo)residual can be determined.

The temperature dependence of (AL/Lo)thennal could be well represented by (Ref. 21)

(AL/Lo)the,mal = A(T – 293)3 + B(T – 293)2 + QT – 293) + D (2)

where A, B, C and D are constants. Table 1 shows the values of these constants detennined from a

polynomial regression of the experimental data for the three specimens and the corresponding coefficients

of determination, Rd2 , for each thermal cycle. The magnitudes of the constants determined for the three

specimens are similar thereby confirming the reproducibility in the data. The values of (AL/Lo)then„al were

subtracted from the measured values of (AL/Lo)first cycle to determine the temperature dependence of

(AL/Lo)residual (Fig. 2(a) to (c)).

TABLE 1.—VALUES OF THE REGRESSION CONSTANTS FOR VPS Cu-8%Cr COATING

Specimen

no.

Cycle description A

(K-3)

B

(K `)

C

(K-')

D Rd -

1 ls` cool-down to3' cool-clown 4.7X10- –3.3x10- 1.9X10- –3.4x10- 0.994

2 l” cool-do« n to 3' cool-down 5.3x10 - –4.1 x 10- 1.9X 10- –1.7x 10- 0.999

3 1” cool-down to 3' cool-down 5.2x 10- –3.8x 10- 1.9X 10- –1.4x 10- 0.999

Combined data for

specimens 1, 2 and 3

I" cool-down to 3' cool-down 5.1 x 10- –3.7x 10-' 1.9x 10- –2.2x 10- 0.972

Figure 2(d) compares the regressed curves' for the three specimens based on the constants given in

Table 1 with literature data for Cr (Ref. 21) and Cu (Ref. 21). The regressed values for the three

specimens are in excellent agreement and closely follow the trend exhibited by Cu (Ref. 21) while

deviating considerably from the thermal expansion data for Cr (Ref. 21). This is not surprising since Cr

has low solubility in Cu (Ref. 22) so that for practical purposes the thermal expansion of Cu-8%Cr should

be similar to that of Cu given the relatively low volume fraction, V f, of the a-Cr particles (Vf –9%). The

excellent agreement between the predicted thermal expansion for the alloy based on a rule of mixtures

(ROM) model with the experimental data provides additional confirmation (Fig. 2(d)).

2The regressed curves were corrected for the residual strain offset at room temperature to ensure that AL'Lo values

are zero at room temperature. Similar corrections were made for the data for the other coating alloys.

NASA/TM-2010-216233
2.0 2.0

Cu-8%Cr

Specimen 1

..... 1st heat up -/

- - 1st-cool down •'^>

---2nd heat up •• i

- -2nd cool down

-----3rd heat up

- - 3rd cool down

-Regression

-Residual strain

fa)

0 200 400 600 800 1000 1200 1400

T(K)

Cu-B Cr

Specimen 2

.....1 st heat up

- 1st-cool down

---2nd heat up

--2nd cool down

----3rd heat up

- -3rd cool down

-Regression

--Residual strain

(b)

200 400 600 800 1000 1200 1400

T (K)

1.5

1.0

0

0.5 J

d

0.0

-05

2.0

1.5

71.0

J

< 0.5

0.0

Cu-S%Cr

-Specimen 1

----Specimen 2

• - -Specimen 3

---ROM

Cr (Touloukian et al. (1975)}

-Cu (Touloukian etal. (1975))

(d)

1.5

-1.0

0.5

0.0

-0 5

0

3.0

2.5

2.0

1.5

J 1.0

0.5

0.0

-0 5

Cu-8%Cr

Specimen 3

1st heat up

- 1st-cool down

---2nd heat up

--2nd cool down

-----3rd heat up

- - 3rd cool down

-Regression

-Residual Strain

(c)

0.5

0 200 400 600 800 1000 1200 1400 0 200 400 600 800 1000 1200 1400

T (K) T (K)

Figure 2.-Temperature variation of AULo and residual strain for three Cu-8 %Cr specimens during three heat up-cool

down cycles; (a) specimen 1; (b) specimen 2; (c) specimen 3; (d) comparison of the present results with literature

data for Cr and Cu [21] as well as with rule of mixture (ROM) predictions.

3.2.2 Cu-26%Cr

Figure 3(a) shows the temperature dependence of the thermal expansion for Cu-26%Cr. In contrast to

Cu-8%Cr (Fig. 2), the values of AL/Lo for the first heat-up cycle are lower than the measured values for

the first cool-down and other subsequent thermal cycles. Figure 3(b) compares the regressed values after

correcting for the room temperature residual strain offset with the data for Cr and Cu (Ref. 21), as well as

the predicted values from a ROM model (Vf -30%). Equation (2) is an excellent fit to the experimental

data, where the regression parameters and values of Rd2, are given in Table 2. Once again, the differences

between the first heat-up cycle and the subsequent thermal cycles are attributed to residual stresses

developed in the coating during processing. The corresponding temperature dependence of residual strain

is also shown in Figure 3(a). The magnitudes of the uniaxial residual strain are compressive in Cu-26%Cr

compared to the tensile strains observed in Cu-8%Cr (Figs. 2(a) to (c)), which suggests that there is no

simple relationship between the development of residual stresses during processing and the Cr content of

the alloys. A comparison of Figures 2(d) and 3(b) reveals that the thermal expansion of Cu-Cr alloy

coatings decrease with increasing Cr content consistent with the ROM model.

TABLF. 2 -VALTTF.S OF TT4F FTTTTNCT CONSTANTS FOR VPS Cn-7.6%Cr COATTNCr

Specimen

110.

Cycle description A

(KB

(K_^l C

-1)

D Rd1 1't cool-down to 3 Id cool-down 3.8X 10- -1.9x 10-1 1.6x 10- 1.5x 10- 0.999

NASA/TM-2010-216233 5
Cu-8%Cr-1 %AI

- P re s e n t l n ve st i g a t i o n

Cr (Touloukianetal. (1975))

--Cu (Touloukian etas (1975))

(b)

2.0

1.5

1.0

0

J

0.5

c

0.0

-05

2.0

C u-8%Cr-1 %AI

•••••1st heat up

1.5 _-1st-ol down

---2nd heat up

--2nd cool down

0 1.0 ..... 3,d heat up

- - 3rd cool down

-Regression

0.5 -Residual strain

0.0

-0 5 (a)

2.0 2.0

1.5

1.0

0.5

0.0

-05

1.5

1.0

0.5

0.0

-05 (a)

Cu-26%Cr

-1st heat up

1 st-cool down

- - -2nd heat up

--2nd cool down

---- 3rd heat up

- - 3rd cool down

-Regression

-Residual strain

Cu-26%Cr

-Present investigation

. ..ROM

- -Cr (Touloukian et al. (1975))

-Cu (Touloukian et al. (1975))

(b)

0 200 400 600 800 1000 1200 1400 0 200 400 600 800 1000 1200 1400

T (K) T (K)

Figure 3.-(a) Temperature variation of AL/Lc) and residual strain for Cu-26%Cr during three heat up-cool down

cycles. (b) The present results are compared with literature data for Cr and Cu (Ref. 21) as well as with rule of

mixture (ROM) predictions.

0 200 400 600 800 1000 1200 1400 0 200 400 600 800 1000 1200 1400

T (K) T (K)

Figure 4.-(a) Temperature variation of AULO and residual strain for Cu-8%Cr-1 %AI during three heat up-cool down

cycles. (b) The present results are compared with literature data for Cr and Cu (Ref. 21).

3.2.3 Cu-8%Cr-1 %A1

While Cr solubility in the Cu matrix is negligible in the Cu-Cr alloys (Ref. 22), Al is mainly in solid

solution in the Cu matrix in the Cu-Cr-AI system (Refs.23 and 24). Figure 4(a) shows the temperature

dependence of AL/Lo for Cu-8%Cr-1%Al for the first heat-up and subsequent thermal cycles,

respectively. Figure 4(b) compares the present results with literature data for Cr and Cu (Ref. 21). The

regression parameters using Equation (2) and the corresponding Rt(` are given in Table 3. The thermal

expansion values in the first heat-up cycle are almost identical to those measured in the subsequent

thermal cycles indicating that the residual strain is negligible. An examination of Figure 4(b) shows that

the thermal expansion of Cu-8%Cr-1%Al is slightly lower than that for Cu and Cu-8%Cr (Fig. 2(d)).

TABLE 3.-VALUES OF THE FITTING CONSTANTS FOR VPS Cu-8%Cr-1%Al COATING

Specimen

no.

Cycle description A

(g-s

B

K-2)

C

-1

D Ra1 1't cool-down to 3' cool-down 43x 10-10 -L6x 10-7 1.7x 10-3 -4.4 x 10-' 0.998

NASA/TM-2010-216233 6
2.0 2.0

1.5

1.0

0

J

^j 0.5

0.0

-0.5

0

2.0

1.5

1.0

e 0.5

0.0

1.5

1.0

0

0.5

0.0

-0.5 U

0

2.0

1.5

1.0

J

0.5

0.0

NiAI

Specimen 1

.....1 st heat up

- 1st-cooldown

---2nd heat up

--2nd cool down

----3rd heat up

- - 3rd cool down

Regression

-Residual strain

200 400 600 800 1000 1200 1400

T (K)

NIAI

Specimen 3

. ..lst heat up

1st-cool down

- --2nd heat up

--2nd cool down

----. 3rd heat up

- - - 3rd cool down

-Regression

-ReSlUdal strain

NiAI

Specimen 2

•--•1st heat up

- 1st-cool down

---2nd heat up

--2nd cool down

----3rd heat up

- - 3rd cool down

-Regression

-Residual strain

200 400 600 800 1000 1200 1400

T (K)

NiAI

-Specimen 1

--- Speci men 2

- -Specimen 3

-•-Touloukian (1967)

-e-Clark & Whittenherger (1984)

-0.5 1 1-1 1 -0.5 ' '

0 200 400 600 800 1000 1200 1400 0 200 400 600 800 1000 1200 1400

T (K) T (K)

Figure 5.-Temperature variation of AL/Lo and residual strain for three NiAI specimens during three heat up-cool

down cycles; (a) Specimen 1; (b) specimen 2; (c) specimen 3; (d) comparison of the present results with literature

data for NiAI (Refs. 25 and 26).

3.2.4 NiAl

Figures 5(a) to (c) show the temperature dependence of the thermal expansion for three NiAI

specimens. The magnitudes of OL/1 0 in the first heat-up cycle are nearly identical to those determined in

the subsequent thermal cycles similar to the observations on Cu-8%Cr-1%Al (Fig. 4(a)) coatings. Table 4

gives the corresponding regression parameters in Equation (2) and values of Rd for the three NiAI

specimens. These constants are in excellent agreement thereby suggesting that the specimen-to-specimen

variation is negligible (Fig. 5(d)). There are no significant differences between the values of AL/Lo

measured during the first heat-up cycle and the other subsequent cool-down and heat-up cycles

corresponding to negligible residual strains. Literature data on hot-pressed NiAI (Ref. 25) and hotextruded NiAI (Ref. 26) are shown for comparison in Figure 5(d). The present data are in excellent

agreement with the literature data (Refs. 25 and 26), which suggests that the temperature dependence of

thermal expansion of NiAI is not significantly influenced by the processing method.

TABLE 4.-VALUES OF THE REGRESSION CONSTANTS FOR VPS N1Al COATINGS

Specimen

no.

Cycle description A

(K-3)

B

(K72)

C

(K-t)

D Rd'

1 I" cool-down to 3 d 3.1 X 10-"' -2.9X 10-7 X 10- -4.0X 10-- 0.999

2 I" cool-down to 3 d 3.0X 10- -3.OX 10- 1.6X l0- -2.1 X 10-- 0.999

3 1" cool-down to 3' cool-down 3.1 X 10- ' -3.2X 10- 1.6x l0- -2.1 X 10-- 0.999

Combined data for

specimens 1, 2 and 3

I" cool-down to 3' cool-down 3.1 X 10- -3.OX 10- 1.6X 10- -2.7X 10-- 0.998

NASA/TM-2010-216233 7
2.0

1.5

1.0

J

_j

0.0

-0.5

N 1-17%C r-6 % AI-0.5 % Y

•••••1 It heat up

—

1st- .1do-

- - -2nd heat up

— — 2nd cool down

----3rd heel up

- - - 3rd cool do-

-Regression /-

^Residualstrain

—NI-20Crv11Al-0.SY (Taylor& Walsh M04))`'l /

200 400 600 800 1000 1200 1400

T (K)

Figure 6.—Temperature variation of AUL O and residual

strain for Ni-17%Cr-6%AI-0.5%Y during three heat

up-cool down cycles. The present results are

compared with the predicted results for Ni-20%Cr11 %AI-0.5%Y (Ref. 27).

3.2.5 NiCrAIY

Figure 6 shows the temperature dependence of AL/Lo for the Ni- 17%Cr-6%A1-0.5%Y coating and

compares it with published data on Ni-20%Cr-11%A1-0.5%Y (Ref. 27). Table 5 gives the regression

parameters determined from the first to the third cool-down cycle. Once again, the residual strains are

negligible as the values of AL/Lo for all the heat-up and cool-down cycles are similar within the limits of

experimental scatter. As shown in Fi gure 6, the thermal expansion of Ni-17%Cr-6%A1-0.5%Y is slightly

larger than Ni-20%Cr-11%A1-0.5%Y (Ref. 27), where the deviation between the two sets of data

increases with increasing temperature. Despite the relatively large compositional differences between the

two coating alloys, the differences in the values of AL/Lo are sufficiently small to conclude that other

N1CrAlY coatings with similar compositions would have comparable AL/Lo values.

TABLE 5.—VALUES OF THE FITTING CONSTANTS FOR VPS Ni-17%Cr-6%A1-05%Y COATING

Specimen

no.

Cycle description A

(K-3

BK-z C

K-1

D Rj`

1 1 1" cool-down to 3 `1 cool-down 1.0X 10—q -8.6x 10-7 1.7X10-3 -3.4X10-' 0.998

3.3 Implications for Coating GRCop-84 Combustion Liners

N1CrAlY and NiAI have been proposed as protective top coats for GRCop-84 rocket engine

combustion liners for reusable launch vehicles using Cu-Cr or Cu-Cr-AI bond coats (Ref. 14). Figure 7

compares the thermal expansions for Cu-8%Cr, Cu-26%Cr, Cu-8%Cr-1%Al, N1AI, NiCrAIY and

GRCop-84 (Ref. 7). The differences in the magnitudes of therinal expansions between Cu-8%Cr on one

hand, and NiAI and NiCrAIY on the other are relatively large, and increase with increasing temperature.

These observations suggest that using Cu-8%Cr as a bond coat with either NiAI or NiCrAIY as a top coat

on a GRCop-84 substrate is likely to result in the development of large residual thermal stresses when the

rocket engine is first fired.

NASA/TM-2010-216233
Cu-8 % Cr GRCop-84

-- Cu-26%Cr

Cu-8Cr-1AI— --- Cu-8 % Cr-1 % AI s

Cu-8Cr

NiAI ,^° ':'NiCrAIY

--- NI-17%Cr-6%AI-0.5%Y

NiAI

1

Cu-26Cr

2.0

1.5

1.0

J

Q

0.5

0.0

200 400 600 800 1000 1200 1400

T (K)

Figure 7.—Comparative plot showing the variation of

AL/Lo with absolute temperature for VPS Cu-8%Cr, Cu26%Cr, Cu-8%Cr-1%Al, NiAI and Ni-17%Cr-6%AI0.5%Y monolithic coatings. The data for GRCop-84

(Ref. 7) are also shown for comparison.

On close examination, it is evident that the thermal expansion characteristics of NiCrAIY and Cu26%Cr are fairly similar over a wide temperature range, which suggests that the thermal residual stresses

would be small when N1CrAlY is used as a top coat with a Cu-26%Cr bond coat. 3 In contrast, the

differences in the thermal expansions between a NiAI top coat and a Cu-26%Cr bond coat increase with

increasing temperature thereby limiting the use of NiAI as a top coat due to an increased probability of

coating spallation at higher temperatures. Despite the similarities in the magnitudes of the thermal

expansions of Cu-26%Cr and N1CrAlY, it is important to Mote that the thermal expansion between the

GRCop-84 and Cu-26%Cr increases with increasing temperature. Thus, there is a danger that residual

stresses developed at the Cu-26%Ct;% GRCop-84 interface could weaken the strength of the bond and lead

to coating spallation when NiCrAIY and Cu-26%Cr are used as top and bond coats on a GRCop-84 liner.

Figure 7 shows that the thermal expansion characteristics of the Cu-8%Cr-1 %Al coating are almost

identical to that for GRCop-84. The addition of 1%Al to Cu-8%Cr results in a very close match in the

thermal expansions of Cu-8%Cr-1%Al and GRCop-84. Thus, relatively low residual stresses are likely to

develop at the Cu-8%Cr-1%Al/GRCop-84 interface. However, the differences in the thermal expansions

between Cu-8%Cr-1%Al and NiAI and NiCrAIY are much larger especially at the higher temperatures,

which is likely to lead to the development of significant residual stresses at the bond coat-top coat

interfaces.

It is important to note that the magnitude of residual stresses and their effects on coating spallation are

determined only partially by differences in thermal expansion between the top coat, bond coat and the

GRCop-84 substrate. Other factors, such as elastic modulus, ultimate and yield stress, impurities and

microstructures, also influence the magnitudes of the residual stresses and the probability of coating

spallation so that precise determination of the local residual stress distribution is often difficult to

ascertain. Thermal cycling tests conducted on a NiCrAIY-coated GRCop-84 water-cooled specimen using

either a Cu-8%Cr or a Cu-8%Cr-1%Al bond coat under a high heat flux 1­12/O2 combustion flame did not

show any evidence of coating spallation after 40 cycles (Ref. 14). Although the differences in thermal

expansions between either the Cu-8%Cr or a Cu-8%Cr-1%Al alloy bond coat and the NiCrAIY alloy top

coat increases with increasing temperature, heat transfer analyses of heat fluxes similar to those in a

3 It is important to note that in practice, the Cu-26%Cr coating layer would have residual stresses due to spraying on

the substrate prior to the deposition of NiCrAIY top layer, which complicates modeling the residual stresses in a

relatively simple manner.

NASA/TM-2010-216233
rocket engine reveal that for a top coat thickness of about 0.1 mm, the interface temperatures are likely to

be less than 625 K (Ref. 14). Indeed, finite element analyses of the out-of-plane stresses developed during

these thermal cycles reveal that their magnitudes are identical under these high heat flux conditions for

both Cu-8%Cr and Cu-8%Cr-1%A1 bond coats deposited on GRCop-84 substrates using NiCrAIY top

coats (Ref. 14). In this case, the out-of-plane stress at the N1CrAlY/bond coat interface increased from –

580 to –150 MPa in traversing from the top to the bond coat. The fact that coating spallation did not occur

in these high heat flux thermal cycling experiments (Ref. 14) despite the differences in magnitudes of

AL/Lo for Cu-8%Cr, Cu-8%Cr-I%Al and NiCrAIY (Fig. 7) suggests that these residual stresses are

accommodated by the plasticity of the bond coats.

Although Figure 7 suggests that Cu-26%Cr could be used as suitable bond coat for N1CrAlY top coat,

it is important to note that experimental observations revealed that Cu-26%Cr top coats either spalled in

cyclic oxidation tests (Ref. 28) or blistered under high heat flux conditions (Ref. 14). Unlike the more

ductile NiCrAIY, the brittleness of a NiAI top coat may result in its spallation due to its inability to

accommodate any thermal strain mismatch with the Cu-8%Cr-1%Al bond coat. Therefore, it is concluded

that Cu-8%Cr-1%Al and NiCrAIY are likely to be the most suitable protective bond and top coat

combination for GRCop-84.

Summary and Conclusions

The temperature dependence of the thermal expansion, AL/Lo, of vacuum plasma sprayed Cu-8%Cr,

Cu-26%Cr, Cu-8%Cr- I %Al, NiAI and Ni- 17%Cr-6%Cr-0.5%Al monolithic coatings between 300 and

1223 K are reported. Excellent specimen-to-specimen and cycle-to-cycle reproducibility were observed

for all the coatings except during the first heat-up cycle for which the thermal expansion was different

than those in the subsequent cycles. The observed behavior could be well represented by the equation:

(OL/Lo),h,1,,,a, = A(T – 293) 3 + B(T – 293)2 + C(T – 293) + D. Comparison of the data for Cu-8%Cr and

Cu-26%Cr alloy coatings with literature data for Cr and Cu, as well as rule of mixture predictions,

revealed that the thermal expansion behaviors of these alloys are determined primarily by the Cu-rich

matrix. The data for NiAI are in excellent agreement with published results irrespective of the methods

used for processing the materials. It is concluded that the Cu-8%Cr-1%A1 coating would be the most

suitable bond coat for GRCop-84 with NiCrAIY being the most suitable top coat.

References

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TM X-73665, Lewis Research Center, OH (1977).

2. D.K. Huzel and D.H. Huang, Modern Engineering for Design of Liquid-Propellant Rocket Engines,

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Aeronautics and Astronautics, Inc., Washington, DC, pp. 67-134 (1992).

3. J. Singh, G. Jerman, B. Bhat and R. Poorman, "Microstructural Stability of Wrought, Laser And

Electron Beam Glazed NARloy-Z Alloy At Elevated Temperatures," NASA TM-108431, George C.

Marshall Space Flight Center, Huntsville, AL, (1993).

4. H.J. Kasper, "Thrust Chamber Life Prediction," Advanced High Pressure 0 2 /H2 Technology, NASA

CP-2372 (eds. S.F. Morea and S.T. Wu), George C. Marshall Space Flight Center, Huntsville, AL,

pp. 36-43 (1985).

5. R.J. Quentmeyer, "Rocket Thrust Chamber Thermal Barrier Coatings," Advanced High Pressure

02/H2 Technology, NASA CP-2372 (eds. S.F. Morea and S.T. Wu), George C. Marshall Space Flight

Center, Huntsville, AL, pp. 49-58 (1985).

6. D.B. Morgan and A.C. Kobayashi, " Main Combustion Chamber and Cooling Technology Study –

Final Report," NASA CR-184345, NASA Marshall Space Flight Center, Huntsville, AL (1989).

NASA/TM-2010-216233 10
7. D. Ellis and D. Keller, Thermophysical Properties of GRCop-84, NASA/CR-2000-210055, NASA

Glenn Research Center, Cleveland, OH (2000).

8. R. Holmes, D. Ellis and T. McKechnie, "Robust Low Cost Aerospike;RLV Combustion Chamber by

Advanced Vacuum Plasma Process," Countdown to the Millennium Proceedings, 36 h Space

Congress, Canaveral Council of Technical Societies, Cape Canaveral, FL (1999).

9. R. Hickman, T. McKechnie and R. Holmes, "Material Properties of Vacuum Plasma Sprayed Cu8Cr-4Nb for Liquid Rocket Engines," Paper No. AIAA-2001-3693, 37 1h AIAA/ASME/SAE/

ASEE/Joint Propulsion Conference, Salt Lake City, UT, American Institute of Aeronautics and

Astronautics, Inc., pp. 1-9 (2001).

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13. S. Elam, R. Holmes, T. McKechnie, R. Hickman and T. Pickens, "VPS GRCop-84 Chamber Liner

Development Efforts," 52"d JANNAF Propulsion Meeting/ I" Liquid Propulsion Subcommittee

Meeting, Las Vegas, Chemical Propulsion Information Agency, The Johns Hopkins University,

Baltimore, MD, pp. 1-10 (2004).

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(2007).

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19. S.V. Raj, L.J. Ghosn, A. Agarwal and T.P. Lachtrup, "An Assessment of the Residual Stresses in

Low Pressure Plasma Sprayed Coatings on an Advanced Copper Alloy," Surface Engineering: In

Materials Science II (eds. S. Seal, N.B. Dahotre, J.J. Moore, C. Suryanarayana and A. Agarwal) The

Minerals, Metals & Materials Society, Warrendale, PA, pp. 49-56 (2003).

20. Z. Gan, H.W. Ng and A. Devasenapathi, Surf. Coat. Technol. 187, 307-319 (2004).

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and A71oys, Thermophysical Propertied of Matter, vol. 12, IFI/Plenum, New York, NY, (1975).

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International, Materials Park, OH (1990).

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14. ABSTRACT

Metallic Cu-8%Cr, Cu-26%Cr, Cu-8%Cr-1%Al, NiAl and NiCrAIY monolithic coatings were fabricated by vacuum plasma spray

deposition processes for thermal expansion property measurements between 293 and 1223 K. The corrected thernal expansion, (DL/L,)

varies with the absolute temperature, T, as (DL/L0)rhennat — A(T - 293)3 + BIT - 293)2 + C(T - 293) + D, where, A, B, C and D are thermal,

regression constants. Excellent reproducibility was observed for all of the coatings except for data obtained on the Cu-8%Cr and Cu-26%Cr

coatings in the first heat-up cycle, which deviated from those determined in the subsequent cycles. This deviation is attributed to the

presence of residual stresses developed during the spraying of the coatings, which are relieved after the first heat-up cycle. In the cases of

Cu-8%Cr and NiAl, the thermal expansion data were observed to be reproducible for three specimens. The linear expansion data for Cu-8%

Cr and Cu-26%Cr agree extremely well with rule of mixture (ROM) predictions. Comparison of the data for the Cu-8%Cr coating with

literature data for Cr and Cu revealed that the thermal expansion behavior of this alloy is determined by the Cu-rich matrix. The data for

NiAl and NiCrAlY are in excellent agreement with published results irrespective of composition and the methods used for processing the

materials. The im lications of these results on coating GRCo -84 copper allo y combustor liners for reusable launch vehicles are discussed.

15. SUBJECT TERMS

Thermal expansion, Vacuum plasma sprayed coatings; Copper alloy coatings; NiAl coating, NiCr'AlY coating; Launch vehicles

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