FR-19009-1

10 DECEMBER 1985

ADVANCED TURBINE STUDY

FINAL REPORT

Prepared Under Contract NAS8-33821

for

National Aeronautics and Space Administration

George C. Marshall Space Flight Center

Marshall Space Center, Alabama 35812

Prepared by

Pratt & Whitney

Government Products Division

P. O. Box 2691, West Palm Beach, Florida 33402

•"'

'85 N86-20495

21H

•3/20 04272

UNITED

TECHNOLOGIES

PRATT&WHITNEY

27

^ ihe Untied Slates ol
FR-19009-1

10 DECEMBER 1985

ADVANCED TURBINE STUDY

FINAL REPORT

Prepared Under Contract NAS8-33821

for

National Aeronautics and Space Administration

George C. Marshall Space Flight Center

Marshall Space Center, Alabama 35812

Prepared by

Pratt & Whitney

Government Products Division

P. O. Box 2691, West Palm Beach, Florida 33402

UNITED

TECHNOLOGIES

PRATT&WHITNEY

Printed in the United States ol America
PratL& JVhitney

FR-19009-1

FOREWORD

This final report presents the results of the work performed under Contract NAS8-33821,

covering the period 1 August 1982 to 15 August 1985. This work was conducted for the National

Aeronautics and Space Administration George C. Marshall Space Flight Center (NASA MSFC)

by Pratt & Whitney/Government Products Division (P&W/GPD). Mr. J. H. Castro was the

engineering manager for this effort and author of this report. The following personnel have

contributed to the material presented in this report: D. E. Paulus, P. A. Allard, H. M. Gibson,

J. R. Mullaly, and T. L. Woods.
I

Pratt & Whitney

FR-19009-1

CONTENTS

Section Page

I INTRODUCTION 1-1

II SUMMARY IM

HI TECHNICAL ACTIVITY — ORIGINAL PROGRAM III-l

A. Task 1 — Case 2 and 3 Analysis III-l

B. Task 2 — Subtask 201, Specimen Design III-8

C. Task 2 — Subtask 202, Specimen Fabrication IH-17

D. Task 2 — Subtask 203, Specimen Test 111-50

E. Task 3 — Subtask 301, Airfoil Design and Analysis 111-51

F. Task 4 — Airfoil Fabrication and Task 5 — Test Follow Up 111-70

IV TECHNICAL ACTIVITY — REDIRECTED PROGRAM IV-1

A. Task 1 — Hardware Design IV-1

B. Task 2 -r- Hardware Fabrication IV-7

C. Task 3 — Turbine Blade Testing and Data Analysis IV-13

V CONCLUSIONS AND COMMENTS V-l

APPENDIX A. SUMMARY OF SPUTTER DEPOSITION WORK A-l

APPENDIX B. ADVANCED TURBINE STUDY ENGINEERING

DRAWINGS B-l

APPENDIX C. VERIFICATION OF ATS BLADE GRAIN QUALITY AND ORIENTATION BY GRAIN ETCH AND LAUE X-RAY

METHODS C-l

ui
Pratt & Whitney

FR-19009-1

SECTION II

SUMMARY

The original objective of this phase of the Advanced Turbine Study (ATS) Program was to

investigate the feasibility of an advanced convective cooling concept applied to rocket turbine

airfoils operating in a high-pressure hydrogen and methane environment. The concept consisted

of a central structural member in which grooves are machined or etched. The grooves are

temporarily filled with a removable filler and the entire airfoil is covered with a layer of

electroformed or sputtered nickel, or nickel base alloy. After the filler is removed, the low

thermal resistance of the thin nickel closure allows the wall temperature to be significantly

reduced by heat transfer to the coolant.

The program was divided into six tasks. The following describe the six original tasks as

presented by the program plan submitted to NASA on 27 August 1982.

• Task 1 — Turbine Performance Appraisal

The thermal analyses of the Case 2 and Case 3 turbines, as defined in NASA

Contract NAS8-33821 modification S/A-1, will be revised using the advanced convective cooling techniques.

The results will be compared to test data under Task 5 to gage the accuracy

of the analytical techniques. In Case 2, hydrogen is used as the coolant

medium with 80 pounds/second flow. In Case 3, the coolant used is methane

with a 280 pounds/second flow.

• Task 2 — Coolant Geometry Evaluation

Cylindrical test specimens will be used to establish fabrication techniques

and determine the optimum coolant geometry. Cast cylinders of a nickel

alloy will be cored to achieve a wall thickness approximating the airfoil wall

thickness. Analytically chosen coolant geometries will be intalled on the

outer diameter (OD) surface of the cylinders. The cylinders will be subjected

to fluidized bed tests which will thermally cycle each coolant geometry in the

same manner. The data from these tests will be used to select the airfoil

coolant geometries.

• Task 3 — Test Hardware Design and Analysis

Design and analysis of test hardware for verification of a hydrogen and

methane advanced cooling concept will be accomplished under this task. A

common turbine airfoil contour will be used for both hydrogen and methane

testing. The airfoil contour will be a constant cross section from root to tip,

without twist, and designed to fit in the Marshall Space Flight Center

(MSFC) "Space Shuttle Main Engine (SSME) Turbine Airfoil Tester." The

airfoil coolant groove designs will be based on the thermal shock test results

accomplished under Task 2. The groove designs may be different for the

hydrogen and methane test airfoils, if analysis indicates a change is required

for improved durability. A typical airfoil cross section is shown in

Figure H-l.

The coolant feed system to the test airfoil will also be designed under this

task. It will be a mount plate similar to Item 5 on the MSFC test rig drawing,

H-l
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FR-19009-1

SECTION I

INTRODUCTION

In essentially all liquid-fueled rocket engines developed previously, turbine life and inlet

temperature were not major design factors. Since these engines were generally used for

expendable missions, life requirements were comparatively low. The engines were of medium

chamber pressure (i.e., 500 to 1000 psia) and most used the gas generator cycle. Power could be

maintained in the event of any efficiency reduction or excess pressure loss by increasing turbine

flowrate, rather than turbine inlet temperature (TIT), with only a minor effect on overall engine

performance. However, the advent of the reusable, high-pressure staged combustion engine has

completely changed this situation. Since these systems are intended to be reused, the turbine has

to be designed for an engine life of several hundred cycles. Any performance shortcomings in the

engine, due to excessive parasitic flows, high-pressure losses, or low efficiency, can be overcome

only by increasing the TIT. Therefore, turbine technology represents a critical design limit for

the next generation of reusable main engines.

In the first phase of the Advanced Turbine Study (ATS), it was found that conventional

turbine airfoil cooling schemes would not provide adequate cooling durability in a high

temperature and pressure rocket turbine environment (Reference Final Report FR-15978). The

reason for this is that conventionally cooled airfoils are mainly cooled convectively through a

0.04 to 0.06 inch thick wall. This type of cooling in a high-heat-flux, high-temperature rocket

turbine environment creates huge thermal gradients with very little cooling (approximately 50°F)

of the hot side wall. The hot side wall would then experience rapid fatigue cracking and melting

in such an environment.

The current phase of the Advanced Turbine Study investigated an advanced convective

cooling scheme for high-pressure, high-temperature rocket turbine vanes and blades. This

advanced cooling concept consists of a turbine airfoil in which cooling channels (grooves) are

machined into the exterior surface. The grooves are temporarily filled with a removable filler and

the entire airfoil is covered with a thin layer of sputtered nickel alloy. The filler is then removed.

When coolant flows through the channels, the low thermal resistance of the thin, high

conductivity layer allows the hot side wall temperature to be significantly reduced by minimizing

the gradient through the wall. The advantage of this scheme is that the cooling passages are

located close to the hot side wall, thus minimizing the thermal gradient and providing better hot

side wall cooling. The majority of the wall runs very cool and provides adequate strength to carry

gas bending and centrifugal loads.

Due to difficulties encountered during fabrication of the hardware for this cooling scheme,

it was determined that an extensive development program would be required to incorporate the

advanced cooling concept into a turbine blade design. Since such a development program was out

of the scope of the Advanced Turbine Study, the program was redirected by NASA to fabricate

and test a "hot core" type of blade design which will provide improved turbine blade life over a

conventional solid blade. This type of hollow blade, with hot gasses flowing through it, can be

tested in the NASA facility to quantify its life improvement. In P&W's hot core configuration,

the blades are fabricated from PWA 1480 single-crystal material. The enhanced material

properties obtained from this alloy (compared with the directionally solidified alloy currently

used in the SSME) plus the lower strain produced by the hot core configuration, result in a blade

with improved life characteristics. This improved life characteristic can, if necessary, be traded

for an increase in the operating turbine inlet temperature, to provide increased output

horsepower for a given life requirement. The improved life can also be traded for a reduction in

turbine replacement costs, at given inlet temperature and mission requirements.

1-1
Pratt & WhitneyFR-19009-1

Section A-A, and will contain coolant passages and instrumentation access.

Coordination with MSFC to include detailed test rig drawings, and

capabilities will be required to ensure compatible test articles. Figure II-2

shows the proposed fixture.

A test program and required instrumentation will be recommended to

properly evaluate the cooling design and verify the analysis. This test will be

performed by NASA at the MSFC's "SSME Turbine Blade Tester."

Task 4 — Test Airfoil Fabrication

Two cooled airfoils will be fabricated: one for hydrogen operation and one for

methane. In addition, two uncooled airfoils will be provided as slaves. The

cooled airfoils will contain the necessary provisions for pressure and

temperature measurements and will be delivered with the instrumentation

installed.

One mount plate will be fabricated to contain either test airfoil and two

slaves. It will contain a coolant passage from the airfoil attachment to an AN

fitting on the outside plate, and an instrumentation passage. This plate will

attach to the SSME Turbine Blade Tester in the same fashion as the

existing mount.

Task 5 — Analysis of Test Results

The testing of the two articles will be done by NASA in their SSME Turbine

Blade Tester. Instrumentation readings during the testing of the cooling

scheme will ensure that the hot gas and coolant temperature and pressure

match the conditions used in the analysis. Pratt & Whitney will analyze the

metal temperature data, perform metallurgical evaluation of the airfoils, and

adjust the cyclic life prediction to agree with the physical hardware

condition. Similar adjustments to cyclic life predictions will be done for

Cases 2 and 3.

Task 6 — Reports

Bimonthly progress reports will be submitted in the same manner as the

previous Advanced Turbine Study effort. A final report will be issued at the

end of the contract period.

II-2
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FR-19009-1

Coolant Passages

sThin Cover of Electroformed

or Sputtered Nickel Alloy

Figure II-l. Advanced Convection Cooling Scheme

FD 301916

II-3
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FR-19009-1

: Copied Test_Blade

•Uncooled Slave

Airfoil (2)

View B

Instrumentation

Coolant Flow

Test Fixture Similar

to Existing

MSFC Design

-Bolted Coverplate

View B

FD 301917

Figure 11-2. Proposed Test Fixture

The program, initiated in August 1982, remained on schedule throughout the selection of

the candidate materials, the design of the fluidized bed test specimens, the preliminary design of

the test hardware, and the reevaluation of the thermal analysis of Case 2 and Case 3 from the

previous phase of the program. During the fabrication of the fluidized bed test specimens,

difficulties were encountered. These difficulties were caused by poor bond strength between the

test cylinders and the sputtered closeout layer. Several experiments were undertaken to evaluate

and resolve this problem. It was determined that an extensive sputtering developing program

would be required to incorporate the advanced cooling concept into a turbine blade design. Such

a program was out of the scope for the Advanced Turbine Study.

II-4
Pratt & Whitney

FR-19009-1

A different turbine blade concept, a hot core blade fabricated out of PWA 1480 singlecrystal material, was then proposed to NASA by P&W. The hot core design consists of a hollow

blade with an orifice on the pressure side of the blade. Mainstream flow is bled into the core

through this orifice and exits the core at the blade tip. The hot mainstream flow bled into the

core allows the inside of the blade to run hotter, minimizing the blades thermal gradient, and,

therefore, reducing the thermal strain in the blade. The thermal gradients (thermal strains) in

the blades are also reduced by the inherent reduction in wall thickness associated with the hot

core blade. When the enhanced material properties from the PWA 1480 single-crystal material

are used, this reduced thermal strain can be translated into a life improvement of 3.5 times that

of a solid blade design.

The NASA accepted P&W's proposal and the program was redirected to design and

fabricate the hot core blade concept hardware. The hardware was tested by NASA MSFC in their

SSME blade tester rig. The redirected work was accomplished within the original funding bounds

of the program. The redirected program was structured into the following four tasks:

• Task 1 — Hardware Design

Task 1 consists of the design of the hot core blade (Figure II-3) and the

generation of engineering drawings which later were used to fabricate the

blades and blade holding fixture.

• Task 2 — Hardware Fabrication

Task 2 consists of the procurement of raw materials and the fabrication and

instrumentation of the hot core blade hardware.

• Task 3 — Turbine Blade Testing and Data Analysis

Task 3 consists of the performance of the turbine blade tests at NASA and

the evaluation of the test results.

• Task 4 — Reports

Task 4 consists of the bimonthly progress reports and the final report.

The fabrication for the redirected ATS Program hardware was completed on schedule and

within budget. Thirteen blades were delivered to NASA MSFC: nine hollow blades and four solid

blades. A blade holding fixture and associated hardware were also delivered to NASA.

To date MSFC has completed two test series for 71 cycles and 73 cycles each. The general

conclusions obtained from the testing are:

• The preferred P&W thin wall airfoil design with single-crystal material will

meet low cycle fatigue (LCF) life goals

• Initial blade cracking generally occurred on the pressure side, in the platform

fillet radius

• Cored parts are significantly more crack resistant than solid parts

• Thin wall and ventilated parts are the superior configurations for this rocket

environment

II-5
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FR-19009-1

• Based on current testing, it is expected that current high-pressure fuel

turbopump (HPFTP) blade life will be approximately doubled with

PWA 1480 single-crystal material, but will still fall short of reasonable life

goals.

The ATS hardware testing was temporarily halted by MSFC in order to use their test

facility for a different program. The ATS hardware testing will be resumed in early fall 1985.

Four additional blades (two hollow and two solid blades) were delivered to MSFC at

NASA's request under Modification 12 of Contract NAS8-33821. The NASA-supplied material

for these blades was Mar-M 246. This hardware will be tested along with the PWA 1480 singlecrystal blades when NASA resumes testing during fall 1985.

II-6
Pratt & Whitney

FR-19009-1

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SECTION III

TECHNICAL ACTIVITY — ORIGINAL PROGRAM

A. TASK 1 — CASE 2 AND 3 ANALYSIS

The cooling scheme used in the. advanced convective design study for Cases 2 and 3 is

shown in Figure III-l. In this scheme, coolant enters the hollow core and passes radially up into

the blade or vane. From the core, the coolant enters the small chordwise passages through

leading edge holes which connect the core to the passages. The coolant then proceeds chordwise

in the pooves and exits near the trailing edge. For production quantities, the cored blade/vane

and grooves would be integrally casted. The sputtered cover would then have to be applied to the

casting.

FD 267752

Figure III-l. Convectively Cooled Engine Configuration

The coolant grooves are 0.035 inch wide and have 0.045 inch lands between grooves. The

grooves are 0.012 inch deep except for the H2/O2 Ist-stage vanes which have 0.023 inch deep

grooves. The Ist-stage vanes use 2% of turbine inlet flow for cooling and the Ist-stage blades also

use 2% flow for cooling.

1. Task 1 — Subtask 101, Case 2 Analysis

For the case 2 analysis, the fuel used was hydrogen with an 80 pounds/second flow. The

case 2 analysis began with a study to assess the impact of inaccuracies in calculating heat

transfer film coefficients. In any heat transfer analysis the calculation of film coefficients using

empirical correlations is only approximate. An inaccuracy of ± 25% is possible. This parametric

study varied coolant and hot gas heat transfer film coefficients on the Ist-stage turbine blades by

as much as 50%, in order to see the impact on metal temperature and thermal strains. Figures

III-2 through III-4 illustrate the results of this study. Figure III-2 shows that 25% inaccuracy in

film coefficients would result in 10% (or less) inaccuracy in thermal strain range. Figures III-3

and III-4 show that maximum and average metal temperature variation will be less than, or equal

to 75 and 60°F, respectively, for a 25% inaccuracy in film coefficients. These inaccuracies, though

not desirable, are considered acceptable and by assessing their magnitude they can be accounted

for in hardware design.

III-l
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FR-19009-1

0.601—

0.50 —

Strain

Range -

0.40 —

0.30

10 20 30 40

Percent Change in Film Coefficients

FDA 302517

Figure 7/7-2. Case 2 Ist-Stage Blade Strain Range versus Variations in Film Coefficients

III-2
Pratt & Whitney

FR-19009-1

1900

1800

1700

Maximum

Metal

Temperature -

°F

1600

1500

1400

1300

Curve No.

1

2

3

4

5

6

7

8

H

it gas

Const

Const

H

Const

+

Const

10 20 30

Percent Change in Film Coefficients

40 50

FDA 302518

Figure 7/7-3. Cose 2 Ist-Stage Blade Maximum Metal Temperature versus Variations in

Film Coefficients

The turbine horsepower (hp) delivered versus turbine inlet temperature (TIT) curves for

Case 2 (hydrogen) are shown in Figure IH-5. The three labeled curves represent an uncooled

turbine, an advanced convectively cooled turbine using 4% of inlet flow for airfoil cooling, and a

film cooled turbine using 13% of inlet flow for airfoil cooling. The numbered solid lines are lines

of constant cyclic life for the advanced convective cooled turbine, while the numbered dashed

lines are lines of constant cyclic life for. the film cooled turbine.

III-3
Pratt & Whitney

FR-19009-1

900

800

700

Mass

Average

Metal

Temperature

600

Curve No.

1

2

. 3

4

5

6

7

8

Hhot oas

Const

Const

H

Const

+

Const

500 —

400

20 30

Percent Change in Film Coefficients

40 50

FDA 302519

Figure III-4. Case 2 Ist-Stage Blade Average Metal Temperature versus Variations in Film

Coefficients

If life considerations are ignored, the uncooled turbine would show the highest horsepower

output versus TIT with the all convective cooled turbine having the next best horsepower output

and the film cooled turbine showing the lowest horsepower payoff at a given TIT. Table III-l

lists the turbine parameters which influence the horsepower output of the turbine and explains

the differences between the three turbine horsepower levels.

III-4
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FR-19009-1

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Table III-l. Turbine Parameters for Case 2 at 2200°R Turbine Inlet Temperature

Horsepower

W Blade (Ib/sec)*

T Mixed (°R)*

Efficiency (%)

Uncooled

88,200

150

2,200

82.5

Connective Cooled

76,600

139

2,155

77.5

Film Cooled

70,000

134

2,140

75.3

* Mainstream flow and temperature into Ist-stage blade that is available

to do work.

Basically, the less coolant that is used in a turbine, the higher its efficiency and the more

flow there is to do work; consequently, the more horsepower it can deliver.

The horsepower output of the turbine is only one consideration; the other consideration is

turbine life. The comparison of the horsepower output of a cooled versus an uncooled turbine

should be done at a constant life level. The following conclusions are drawn by using the lines of

constant life when comparing overall turbine capability.

As shown in Figure III-5, the advanced convective cooled turbine offers an advantage over

the film cooled turbine. For example, for a 2400 cycle life, the advanced convectively cooled

turbine can obtain 80,000 horsepower at 2315°R TIT, which is a 4% gain over the uncooled

turbine. For the same life, the film cooled turbine could obtain 74,000 horsepower at 2350°R TIT

which is a 4% decrease over the uncooled turbine.

For the high TIT levels (2400°R) and for lower cyclic life, the advanced convectively cooled

turbine shows no horsepower gain over the uncooled turbine. It should be noted that the

uncooled turbine blades were optimized in this study to reduce transient thermal strains and thus

should provide greater life than the current space shuttle hardware, and that an optimized

advanced cooled turbine could give a greater than 4% performance gain over current space

shuttle main engine (SSME) hardware.

The low cycle fatigue (LCF) life curves shown in Figure III-5 were calculated using an

equation which is a function of strain range and metal temperature. The initial study (i.e., first

phase of the program) used a curve of LCF life versus strain range which was not temperature

dependent. This modification will make the LCF lives in Figure III-5 more accurate.

2. Task 1 — Subtask 102, Case 3 Analysis Methane

In Case 3, the fuel used was methane with a 280 pounds/second flow. The turbine

horsepower delivered versus TIT curves for Case 3 (methane) are shown in Figure III-6. The

three labeled curves are for an uncooled turbine, an advanced convectively cooled turbine using

4% of inlet flow for airfoil cooling, and a film cooled turbine using 12% of inlet flow for airfoil

cooling.

If life considerations are ignored, the uncooled turbine would show the highest horsepower

output versus TIT, with the all convective cooled turbine having the next best horsepower output

and the film cooled turbine showing the lowest horsepower payoff at a given TIT. Table III-2

lists the turbine parameters which influence the horsepower output of the turbine and explains

the differences between the three turbine horsepower levels.

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Pratt & Whitney

FR-19009-1

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Table 111-2. Turbine Parameters for Case 3 at 2600°R Turbine Inlet Temperature

Horsepower

W Blade (lb/sec)*

T Mixed (°R)*

Efficiency (%)

Uncooled

102,100

505

2,600

85.5

Connective Cooled

86,740

466

2,551

80.5

Film Cooled

80,130

453

2,454

79.0

'Mainstream flow and temperature into Ist-stage blade that is available

to do work.

Basically, the less coolant that is used in a turbine the higher its efficiency and the more

flow there is to do work; consequently, the more horsepower it can deliver.

The horsepower output of the turbine is only one consideration; the other consideration is

turbine life. The comparison of the horsepower output of a cooled versus an uncooled turbine

should be done at a constant life level. The following conclusions are drawn by using the lines of

constant life when comparing turbine capability.

The payoff for cooling is much greater with the methane cycle than for the hydrogen cycle.

For example, for a 1200 cycle life, the advanced convectively cooled turbine can deliver 69,000

horsepower at 2440°R TIT, and the film cooled turbine delivers 52,500 horsepower. This is a 31%

increase over an uncooled blade with convective cooling and a 28% increase over an uncooled

blade with film cooling. At 2500"R TIT, the film cooled turbine offers more advantage than the

convectively cooled turbine.

The outside film coefficients used in the life analysis of the methane turbine presented

above, were modeled using frozen combustion products for methane. These should provide a

more accurate assessment of the turbine environment. This change accounts for the differences

in the shape of the lines of constant life on the horsepower versus TIT curve, when compared to

the information presented in the final report of the first phase of the Advanced Turbine Study

(FR-15978, dated 12 April 1982, Figure 2.5-3).

B. TASK 2 — SUBTASK 201, SPECIMEN DESIGN

The advanced convective cooling concept investigated in this program is specially suited for

high-pressure, high-temperature rocket turbine vanes and blades. This concept consists of a

turbine airfoil in which cooling channels (grooves) are machined into the exterior surface. The

grooves are temporarily filled with a removable filler and the entire airfoil is covered with a thin

layer of electroplated or sputtered nickel alloy. The filler is then removed. The low thermal

resistance of the thin, high conductivity layer allows the hot side wall temperature to be

significantly reduced by minimizing the gradient through the wall. A typical cross section of a

cooled blade wall is shown in Figure III-7. The advantage of this scheme is that the cooling

passages are located close to the hot side wall, thus minimizing the thermal gradient and

providing better hot side wall cooling. This causes the majority of the wall to run very cool and

provide adequate strength to carry gas bending and centrifugal loads.

III-8
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Unsupported ^

""" Slot Length ""

^ ^

Coolant

Slot

Hot Gas

r

0.010 in. 1 — Sputtered

/ on Cover

' \

._ Land _

Width

' f ' 1 Coolant

Slot

|\

Slot

Depth

\

Core

FDA 301918

Figure 7/7-7. Typica/ Section of Cooled Blade Wall

Early in the program, preference was given to electroplating over sputtering as a method of

fabricating the thin conductive layer around the airfoil. The reasons for this were the simplicity

and the lower cost of electroplating as compared to sputtering.

It was determined that an advanced turbine blade alloy would have to be used as a coolantgroove cover layer, since analysis snowed that pure nickel (the original choice for the cover layer

material) did not have the necessary strength at the required temperatures to withstand the

operating loads (See the Material Selection, Section B.I for details). This rendered electroplating

unacceptable. The limitation of electroplating is that it can only be used to plate pure elements

or elements that are very close in the galvanic series scale (like cobalt and nickel).

The chosen method of deposition of the cover layer is sputtering. The sputtering technique

is a high energy deposition which provides a chemical bond with the bonding lands on the blade.

This process has been used successfully to deposit different alloys on turbine airfoils.

Sputter deposition is accomplished with a system comprised of vacuum pumps and gages,

power supplies, and specifically designed electrodes; all of these parts are collectively referred to

as "deposition geometry." For this program, the deposition geometry shown in Figure III-8 was

used. The parts were rotated in coaxial alignment with a cylindrical target in the water-cooled

sleeve. Solenoid coils surround the deposition geometry and its vacuum housing to provide a

magnetic field, which provides an additional control on ion-current density in the system.

A clean part is installed in the system and the system pumped to ultrahigh vacuum (less

than 10~6 torr). A rare gas, krypton for this program, is bled into the system to bring system

pressure to the 10~3 torr range. A gas flow at constant pressure is maintained with a leak valve

and throttled vacuum pumps. Several modes of gaseous electrical discharge can be established to

effect sputter cleaning, rapid deposition rate, etc., as required. For example, in the triode mode a

plasma is established in the region between the filaments and anode. When a high negative

voltage is applied to the target, ions are accelerated from the plasma. The accelerated ion has

sufficient energy to eject or "sputter" atoms from the target surface. The ejected atoms travel at

high speeds and impact the substrate surface where they accumulate to buildup a cover layer.

The substrate receives a considerable amount of electron bombardment in the triode mode,

which causes the part to heatup. This heatup can be minimized by operating in an inverted

III-9
Pratt & Whitney

FR-19009-1

elevated temperature oxygen and water vapor environment and lack of experience in

manufacturing hollow airfoil shapes with this material. The search for an acceptable material

was switched to the typical cast alloys used in turbine airfoils.

1001—

Coolant to

Mainstream A P

Stress - ksi

90

80

70

60

50

40

30

20

10

Mainstream • •0.010 in.

Max A P Stress

Coolant Passage

Unsupported

Section Through

 Len9th

Blade Wall

P in. = 1.5 P CA t hot gas

AP = 2500 psia

P in. = 1.11 P

CA t hot gas

AP = 1500 psia

I I

10 20 30 40 50 60

Unsupported Length - mils

70 80

FDA 301919

Figure 7/7-9. Stress in Cover Due to Coolant to Mainstream Pressure Differential in Rig

III-ll

PAGE BtW^< SOT F.U8ED
Pratt & Whitney

FR-19009-1

The final selection of the material was based on NASA Report TND-8071, October 1975,

"Comparative Thermal Fatigue Resistance of Twenty-Six Nickel and Cobalt-Base Alloys." The

alloys were thermally shocked in a fluidized bed. A summary of the results is provided in

Figure 111-10. Since the sputtering process of applying the cover gives an equiaxed grain

structure, the material selection for the cover was limited to equiaxed materials. Since the strains

in the remainder of the airfoil are lower than in the cover layer, as shown in Figure III-ll, it is

reasonable to also select an equiaxed material for the airfoil. Also, a better bond might result by

using the same material and grain structure for both the cover layer and the airfoil.

The best equiaxed material was B1900 + Hf (Reference Figure 111-10). However, P&W

experience with cast, production airfoils made of B1900 + Hf is that shrinkage porosity is

encountered in cluster areas. This causes material weaknesses which severely degrade the

material properties. Thus, instead of B1900 + Hf, a similar alloy Mar-M 247 (PWA 1447) was

selected for the ATS blade and cover layer materials. A comparison of the chemical compositions

of the two nickel alloys is presented in Figure 111-12. The castability of Mar-M 247 in solid and

hollow parts is excellent and clustered porosity is not evident. The alloy Mar-M 247 has equal to,

or superior, thermomechanical fatigue properties when compared to B1900 + Hf. Thus, P&W

feels that Mar-M 247 will provide a distinct advantage over B1900 + Hf for the ATS program.

An attempt was made to procure and test two sets of fluidized bed test samples (i.e., the two

top materials identified in the study). Quotes were requested from vendors, and the impact to the

program was investigated. The investigation showed that it would have been too costly and out of

the program scope. Also, because of the lead times involved for the material procurement and

manufacturing of the blades for the SSME Blade Tester, a decision on the blade material had to

be made before the fluidized bed test results became available. For these reasons, it was decided

that it was not feasible to pursue the two top materials; the decision was made to continue the

program within its original scope using Mar-M 247 (PWA 1447), the best material choice.

2. Coolant Geometry Determination

In order to select the optimum coolant groove-land geometry combination, analytically

chosen coolant geometries would be machined on the outer diameter (OD) surfaces of cast test

cylinders. The grooves would then be filled and the cylinders covered with the thin conductive

nickel alloy layer. Upon removing the filler material, the cylinder would be subjected to thermal

cycling in a fluidized bed test. These test results would then be used to select the optimum

coolant groove-land geometry combination.

To help select the coolant groove-land geometry, an analytical study was conducted at

elevated temperature engine conditions. The study varied the distances between coolant passages

(land widths) and determined the maximum metal temperature. The results of the study are

shown in Figure HI-13. Land width must be minimized by selecting the smallest land that can

provide a structurally adequate bond plane. A small land will minimize the maximum metal

temperature encountered on the foil hot surface.

A separate study was also performed to evaluate the effects of coolant-to-mainstream

pressure differential on coolant groove width. As shown in Figure III-9, large groove widths will

result in very large stresses in the blade cover layer.

111-12
Pratt & Whitney

FR-19009-1

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FR-19009-1

Cover

0.010 in.

Thick

I

-0.36%

(1325)

-0.18%

(1065)

-0.11%

(975)

Coolant Slot

+0.3%

(290)

1

+0.2%

(380)

0.1%

(440)

Hot Gas

r

-0.35%

(1340) r

-0.35%

(1380)

-0.17%

(1100)

-0.21%

(1180)

+0.13%

(650)

+0.2%

(480)

+0.08%

(725)

-0.03%

(910)

+0.16%

(545)

+0.07%

(705)

(Metal Temperature °F) +0.1%

(595)

+0.1%

(540)

T = 1,700°F

hot gas

T = 70°F

coolant

H = 10.500 Btu/hr ft2 °F

hot gas

H = 4,500

coolant

MAR-M 247 Material

H + O Combustion Products

FDA 301921

Figure III-11. Thermal Strains in Cooled Blade at Steady-State Rig Condition

HI-14
Pratt & Whitney

FR-19009-1

Mar-M 247

B1900 + Hf

Ni

R

R

R

Cr

8.4

8.0

Co

10.0

10.0

Ti

1.05

1.0

Al

5.5

6.0

Mo

0.65

6.0

C

0.15

0.11

W

10.0

0

Ta

3.05

4.25

Hf

1.4

1.15

= Remainder

FDA 301922

Figure 111-12. Chemical Composition Comparison

2300

2250

2200

Maximum Metal

Temperature - °F

2100

2050

— T = 2400 °F

hot gas

T = 450 °F

coolant

H = 30,000 Btu/hr ft2 °F (H + OJ

hot gas ' ,

 x 2 Z

H = 33,000 Btu/hr ft2 °F (HJ

coolant ' v z7

Cover Thickness = 0.010 in.

Conductivity of PWA 1422

0.05

Land Width - in.

0.10

FDA 301924

Figure 111-13. Maximum Metal Temperature versus Land Width for Advanced Engine

Conditions, H2 + 02 Propellants

Based on these studies, six coolant groove-land geometry combinations were selected to be

tested in the fluidized bed test. The groove patterns chosen covered the useful range of groove

depth, groove width, and land width combinations. The maximum groove width was set at 0.05

inch. Larger groove widths would result in very high stresses due to the coolant-to-mainstream

pressure differential. Two groove depths were selected, a shallow groove (0.011 to 0.012 inch) and

a deeper one (0.022 to 0.025 inch). Very shallow grooves would be impractical, since

manufacturing tolerances would cause wide variations in coolant flow area. Very deep grooves

would be impractical because they would seriously reduce the load carrying cross sectional area of

the airfoil. Land widths of 0.030, 0.045, and 0.060 inches were selected. Smaller land widths

would provide too little space for a good bond with the sputtered cover. Larger land widths would

reduce the cooling in the center of the land sections due to the longer conduction paths to the

coolant passage.

Figure 111-14 shows the six selected coolant groove land geometry combinations and

illustrates the fluidized bed test specimen assembly. The coolant grooves run perpendicular to

the test cylinder centerline and around its circumference. Each groove has six equally spaced

holes of the same diameter as the groove width. These holes will be used to leach the filler

material after the closeout layer has been applied to the cylinder.

111-15
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Pratt & Whitney

FR-19009-1

Around Cylinder

Figure IH-14. Advanced Turbine Study Fluidized Bed Specimen

FD 253258

As shown in Figure III-14, the coolant grooves run circumferentially (chordwise in an

airfoil) around the test cylinder rather than longitudinally (radially in an airfoil) as originally

proposed. This was the result of a study conducted in which four coolant groove paths were

evaluated. The different groove paths, their advantages and disadvantages are illustrated in

Figure 111-15.

111-16
Trailing Edge

Discharge

Chordwise Grooves

Advantages

1. Rupture of Local Cover Does Not Affect All

Coolant Passages.

2. Coolant Dump Pressure Known at Trailing

Edge.

3. Coolant Feed Path Similar to Existing

Successful Designs.

Disadvantages

1. Grooves Are Perpendicular to Pull Stress.

Deep Grooves Would Reduce Pull Load

Carrying Area. Manufacturing Limitations Will

Be Determined for Groove Depth.

Tip

Discharge -w

Pratt & Whitney

FR-19009-1

Coolant

Feed

Plenum

Tip

Discharge _ \

Radial Grooves With ID Plenum Feed

Advantages

1. Grooves in Direction of Pull Stress.

2. Need Only a Few Coolant Feed Holes in High

Stress ID Region.

Disadvantages

1. Coolant Dump Pressure Not Accurately Known

2. Rupture of Cover Over ID Plenum Would Rob

All Radial Passages.

3. Large Plenum at ID Would Minimize Pull Load

Carrying Area.

Tip Discharge -v

Radial Grooves

Advantages

1. Grooves in Direction of Pull Stress.

2. Rupture of Any Cover Does Not Affect the

Rest of Passages.

Disadvantages

1. Multi-Feed Holes in High Stress ID Region

2. Coolant Dump Pressure Not Accurately Known.

3. Trailing Edge May Be Difficult To Cool.

Figure I I I-15. Groove Paths Considered

Spiral Groove

Advantages

1. High Velocity, High Coolant Film Coefficient

Due to Small Coolant Flow Area.

2. Need for Only One Coolant Feed Hole in High

Stress ID Region.

Disadvantages

1. Rupture of Cover Anywhere Robs All Coolant

Downstream.

2. Coolant Groove Depth Has To Be Large in

Order to Keep Coolant Pressure Drop

Reasonable.

3. Coolant Dump Pressure Not Accurately Known.

FD 301925

C. TASK 2 — SUBTASK 202, SPECIMEN FABRICATION

Six cast PWA 1447 fluidized bed test cylinders were procured from the Howmet Dover

casting facilities. The cylinders were approximately 3 inches long by 2 inches in diameter and

hollow with a wall thickness of approximately 0.200 inch. Upon being received at P&W, the

cylinders were machined to 1.950 ± 0.001 inch outside diameter and a wall thickness of 0.200 ±

0.005 inch.

111-17
Pratt & Whitney

FR-19009-1

On a production basis the coolant grooves would be cast as part of the casting process of the

turbine airfoil. Cast-in grooves should be of high quality and cost effective for larger quantities,

but the cost of die making made it prohibitive for this project.

Because of these reasons, an investigation was undertaken to determine the optimum

method of incorporating the grooves in the fluidized bed test cylinders and the blades. The

investigation results indicated that machining was the optimum way of incorporating the grooves

in the cylinders. Three different methods of incorporating the grooves were investigated: two

chemical milling methods using different masks and a machining method. These methods are

outlined below.

• Chemically Milled Groove No. 1

The groove was chemically milled using PS 249 nickel-base chemical milling

solution at 130°F. The maskant used was Turcoform 522 (PMC 1779)

chemical milling maskant. The maskant was applied by brushing three

separate thin coats on the cylinder and drying between each application. The

groove was made manually by carefully cutting the maskant away with a

razor blade. Although the maskant was removed fairly uniformly, when the

piece was placed in the chemical milling solution the maskant loosened in

various areas and produced a jagged channel 0.024 inch deep in 150 minutes

(Figures 111-16 and 111-17).

• Chemically Milled Groove No. 2

The groove was chemically milled, using PS 249 chemical milling solution

and Tuffil epoxy as the maskant. The Tuffil, which is the consistency of

putty, was very difficult to apply in a uniformly thin manner. The uniformly

thin layer was desired in order to keep the eccentricity to a minimum after

using an 0.020 inch wide cutter to machine the Tuffil from the base material.

The piece then was chemically milled for 150 minutes, producing a groove

0.028 to 0.055 inch wide and 0.013 to 0.025 inch deep. Next, the excess Tuffil

was machined away and the piece was immersed in a heavy duty 20%

potassium hydroxide solution at 200°F to remove the remainder

(Figures 111-18 and 111-19).

• Machined Groove

The first groove was machined using the Monarch NC Machining Center.

The cutter was a solid carbide four-toothed side-cutting milling cutter

0.050 inch thick, intended to produce an 0.050 inch wide channel. The first

pass, 0.011 inch deep, was machined at 1182 rpm and 1.5 inches per minute

feed. This produced no appreciable wear on the cutter, so another 0.011 inch

deep identical pass was made into the same groove. (This resulted in a total

depth of channel of 0.022 inch). Another pass was then made into the same

groove using a feed of 2 inches per minute. Again, no appreciable wear on the

cutter was noted. The last pass was made using a feed of 3 inches per minute

without appreciable cutter wear. The final channel was approximately

0.050 inch wide and 0.040 inch deep. The NC tape was prepared by the

Computer Aided Machining (CAM) Group. No significant difficulties were

encountered machining the alloy, provided that the cutter was solid carbide

(Figures IIM9 and 111-20).

This machining method was chosen because it provided the best dimensional

111-18
ORIGINAL PAGE IS

OF POOR QUALITY

Pratt & Whitney

FR-19009-1

control for the grooves as well as the best groove surface finish, as shown in

Table III-3. The comparison of the machined groove to the chemically milled

Groove No. 2 is shown in Figure 111-21.

After determining the optimum means of incorporating the grooves into the

test cylinders, the remaining cylinders were sent to the shop to be machined,

as shown in Figure HI-14.

FAL 72575

Figure III-16. Chemically Milled Groove No. 1 Detail (10X)

Chemical Milled

Groove No. 1

Chemical Milled

Groove No. 2- Machined

Groove

FD 301932

Figure III-17. Fluidized Bed Test Cylinder — Chemically Milled Groove No. 1 Detail

HI-19
ORIGINAL PAGE IS

OF POOR QUALITY

Pratt & Whitney

FR-19009-1

FAL 72453

Figure 111-18. Chemically Milled Groove No. 2 Detail (10X)

Chemical Milled

Groove No. 1

Chemical Milled

Groove No. 2

Machined

Groove

FD 301933

Figure 111-19. Fluidized Bed Test Cylinder — Chemically Milled Groove No. 2 and

Machined Groove Detail

The finished machined cylinder is shown in Figures 111-22 and 111-23. The machining of the

cylinders was performed using a Monarch NC Machining Center, with an NC tape prepared by

the Computer Aided Machining group. Blade cutters 0.020, 0.035, and 0.050 inch thick were tipground in-house to provide the necessary tools to cut the bottom of the grooves with the required

0.010 inch fillet radius (reference Figure 111-14). The cutting of the grooves proved to be very

difficult and time consuming, especially for the 0.020 inch wide grooves. The difficulty was

caused by the small grooves dimensions and the tight tolerances required (±0.0015 inch). Due to

the complexity of the job, the machining of the cylinders took longer than estimated. Because of

111-20
Pratt & Whitney

FR-19009-1

budgetary reasons, it was decided to cut the number of cylinders to be manufactured from 6 to 4

and eliminate the two sets of 0.020 inch wide slots. The 0.020 inch wide slots were the most

difficult to machine and the most time consuming. In addition, experiments performed in the

laboratory showed that the 0.020 inch wide slots were also the most difficult to fill. Upon

completion of the machining, the cylinders were delivered to the laboratory for the slot filling

and sputtering operations.

FAL 72455

Figure I 11-20. Machined Groove Detail (10X)

Table 7/7-3. Groove Dimensions Comparison in Fluidized Bed Test Cylinder for Machined

Groove and Chemically Milled No. 2 Groove

Location

(deg)

0 (top)

45

90

135

180

225

270

315

Groove

Machined Groove

(in.)

0.052

0.052

0.052

0.052

0.052

0.052

0.053

0.052

Width

Chem Milled

No. 2 Groove

(in.)

0.047

0.055

0.042

0.039

0.034

0.032

0.028

0.034

Groove Depth

Chem Milled

Machined Groove No. 2 Groove

(in.)

0.040

0.040

0.041

0.040

0.040

0.040

0.041

0.039

(in.)

0.025

0.025

0.021

0.018

0.015

0.015

0.013

0.015

It is important to point out that the cylinders could have been machined using a lathe, in a

much more simple and inexpensive manner. But this process would not have been applicable to

machining the blades for the rig, since the blades can not be turned in a lathe. The intent of the

cylinder fabrication was to duplicate the fabrication process to be used in the blades and to learn

from this experience.

111-21
ORIGINAL PAGE IS

OF POOR QUALFTY

Pratt & Whitney

FR-19009-1

FAL 72452

Figure 111-21. Comparison of Chemically Milled Groove No. 2 and Machined Groove (10X)

FE 355156

Figure 7/7-22. Finished Machined Cylinder

111-22
Pratt & Whitney

FR-19009-1

ORIGINAL PAGE IS

JUALFTY

FE 355157

Figure 7/7-23. Finished Machined Cylinder

While the cylinders were being machined, laboratory experiments were underway to obtain

the best material and method to be used in filling the grooves. Prior to sputtering, three methods

of filling the grooves were investigated by using available (this accounts for the different material

and diameter) bar stock in which grooves, which closely approximated the test cylinder groove

dimensions, were machined by using a lathe. These methods were:

• Aluminum wire was pressed into the groove and made to conform to the

groove geometry

• Aluminum was sputtered on the bar to fill the grooves

• A silica base slurry was used to fill the groove.

1. Groove Filling Experiments

a. Aluminum Wire

Grooves of the same approximate size as those in the test cylinders were machined (using a

lathe) into a 0.625 inch diameter section of PWA 658 (IN 100) round bar stock. Aluminum wire

(of unknown purity) was used to fill the grooves. A 0.060 inch diameter wire was pounded (lightly

beaten into place with a small hammer) into the 0.050 inch wide by 0.024 inch deep groove

without difficulty. The wire would not stay in place within the 0.050 inch wide by 0.012 inch deep

groove, probably due to lack of side wall area for effective mechanical holding. The 0.060 inch

diameter aluminum wire was thinned down to 0.040 inch diameter for filling the 0.035 inch wide

grooves and to 0.025 inch diameter for filling the 0.020 inch wide grooves; then, the wire was

installed in the grooves by the method described above.

111-23
Pratt & Whitney

FR-19009-1

Immersion of the aluminum wire into a 7 molar solution of NaOH proved an effective

method of thinning it to any required diameter.

After the grooves were filled with the aluminum wire, the excess aluminum was machined

flush with the round bar surface. The whole specimen was then lightly cleaned using 600 grit

silicon carbide abrasive paper. Figure 111-24 shows the appearance of this specimen after groove

filling.

0.012 in. Deep

0.020 in. Wide

0.024 in. Deep

0.020 in. Wide

0.012 in. Deep

0.035 in. Wide

0.024 in. Deep

0.035 in. Wide

0.012 in. Deep

0.050 in. Wide

0.024 in. Deep

0.050 in. Wide

Magnification 3x

FD 301926

Figure III-24. Aluminum Wires Beat Into Grooves in PWA 658 Round Bar Stock

(Trimmed Back to the Original Surface OD by Lathe and Hand)

b. Sputtered Aluminum

A second specimen, a PWA 1422 round bar of 0.5 inch diameter, was machined on a metal

lathe to produce grooves with the same dimensions as those on the actual cylindrical test

specimen. This grooved rod was then coated with sputtered aluminum in a hollow-cathode DC

coater.

The substrate temperature was kept low, approximately 950°F to avoid melting the

aluminum deposit. The coating cycle was allowed to run for 24 hours, which produced an average

coating thickness of 31.8 mils. The deposit had a lustrous finish, defect-free to the unaided eye.

After removal from the coater, the substrate was fixtured on the 6 inch lathe and turned

back to its original OD, final finishing to remove all aluminum from the OD surface was

111-24
ORIGINAL PAGE IS

OF POOR QUALITY

Pratt & Whitney

FR-19009-1

performed using No. 600 grit silicon carbide abrasive paper. The grooves were examined by using

a photographic technique to obtain profiles of the sputtered deposits, which were used as models

to illustrate the effects of groove dimensions on the filling process. It was observed that the ideal

groove from a sputtered deposit viewpoint would be both wide and shallow. In this case, the

grooves which filled best were 0.050 inch and 0.035 inch wide by 0.012 inch deep. Figure 111-25

shows the PWA 1422 bar before and after groove filling by sputtering and the cross section of the

filled-in groove.

PWA 1422

Grooved Rig Bar

Prior to Coating

With Aluminum

Aluminum Deposit Profiles

0.012 in. Deep

0.020 in. Wide

0.024 in. Deep

0.020 in. Wide

0.012 in. Deep

0.035 in. Wide

0.024 in. Deep

0.035 in. Wide

0.012 in. Deep

0.050 in. Wide

0.024 in. Deep

0.050 in. Wide

Aluminum Deposit

Machined Back

to Original Rig

Bar Surface

Mag 3x

Run 5-83-002

Sputtered Aluminum on

PWA 1422 Rig Bar

As Coated

Mag 3X

Figure I I 1-25. Sputter-Filled Groove in PWA 1422 Bar

111-25

FD 301927
Pratt & Whitney

FR-19009-1

c. Silica Base Slurry

A third specimen was machined in the lathe out of 0.625 inch diameter PWA 658 round bar

stock. A slurry of high porosity fused silicon with a colloidal silica (30% Si02) binder was forced

into the grooves machined in the specimen. After air drying for one hour, the specimen was

heated for 30 minutes to drive off excess moisture, then baked at 600°F for one hour.

Examination of the cured filler material in the grooves showed no separation between the

filler and the groove side wall and no serious shrinkage related problems in the filler.

Figure 111-26 shows the PWA 658 grooved specimen in stages of filling. Figure 111-27 illustrates

one of the grooves filled with the silica and cured for one hour at 600°F.

2. NaOH Leaching of Aluminum (99.99%) and Silica Fillers

The criteria for a successful filler material indicates that it not only has to be applied to fill

the grooves, but it also must be removed after the sputtering operation has taken place. With this

in mind, it was decided to compare the teachability (ability to remove the material by chemical

means) of the two materials in question — aluminum and silica base slurry. The results are

contained in Table III-4.

The leaching rate of the silica material was determined using a 7 molar NaOH solution with

and without ultrasonic cavitation and was compared to the leaching rate of pure aluminum under

the same conditions. Figure 111-28 is a graph illustrating the relative rates of material removal. It

should be noted, however, that this data has been obtained from bulk material which is fully

accessible to the NaOH solution. In actual practice, it is expected that the surface area of the

filler material actually exposed to the leaching solution will be greatly restricted so that the

actual time required to remove a like amount of material may be several orders of magnitude

higher.

Based on the faster leaching rates and the more uniform filling coverage in the cylinder's

grooves, it was decided to use the silica base slurry as the filling material for the grooves in the

fluidized test cylinders and the blades. The one drawback associated with the slurry is that once

it is dry, the cylinders must be handled with great care due to the fragility of the cured slurry.

3. Sputtering the Cover

During subsequent handling of the silica-filled grooved specimen, shown in Figure 111-26,

much of the filler material was jarred loose and lost. With the purpose of obtaining early insight

into the effects of sputtering a metallic overlay on the ceramic filler material, this specimen was

used as a substrate during initial sputtering cycles required to clean up the new PWA 1447

hollow-cathode targets. The first overlay deposit put on this specimen was sputtered on using

magnetic coils around the target OD (outside the chamber walls) which helped maintain a DC

diode discharge near the target inner diameter (ID). This, in effect, kept the plasma from heating

up the substrate. The coating produced by this scheme is shown in Figure 111-29. The deposit had

a coarse, cauliflower-like surface, which indicates an open, columnar structure. The grain size

over the ceramic filler was several times larger than that over the metal. No bonding (deposit

continuity) was shown between the cover on the ceramic and the metal. Any cracks or gaps

originally present on the substrate surface were still evident.

111-26
ORIGINAL PAGE IS

OF POOR QUALITY

Pratt & Whitney

FR-19009-1

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111-28
Pratt & Whitney

FR-19009-1

Table III-4. Leachability Test Results (Weights Shown Are in Grams)

NaOH Solution: 28 grams of pelletized NaOH in 100 ml H20

• 7 molar solution

• Heated to 140 to 160"F

• Static test done on hotplate

• Ultrasonic test done in heated ultrasonic tank

No

Ultra

Sonic

Before

Test

Wt of Si02 0.13596

After After

5 min 10 min

0.0758 0.0236

After

15 min

All material in

solution at

end of 15 min

Wt of Al 0.1265 0.1106 0.0811 0.0578

With Wt of Si02 0.1384 0.0289 Fully DisUltra solved after

Sonic 2.5 min of

this period

Wt of Al 0.12196 0.0909 0.090 0.0696

Alum, (U.Sonic),

NaOH

Alum,

(No U.Sonic),

X NaOH

0.121916 0.1384 gm

U.S.

(No U.Sonic),

NaOH

Si02,

.Sonic

NaOH

No U.S.

10 15

Time (minutes)

0.1265 0.13596 gm

Test Specimens Prior to NaOH Leaching

FD 253290

Figure 7/7-28. Relative Rates of Material Removal

This specimen was left in place without change to be used again as a substrate during a

sputtering cycle which employed a DC triode discharge. In this mode, the plasma is not confined

to the area near the surface of the target and can affect the substrate by ionic bombardment. The

surface of the substrate itself can be sputtered off even as material sputtered from the target is

being deposited. By manipulating the negative bias applied to the substrate, the factors of heat

and back-sputtering can be employed to tailor the coating microstructure within certain limits.

Figure 111-30 shows the effect of using bias during a DC triode sputtering operation. Although the

111-29
Pratt & Whitney

FR-19009-1

original surface, as shown in Figure 111-29, was severely cracked and coarse, this surface shows

the coating over the boundary between the ceramic and metal to be almost fully continuous.

Thus, this was the method employed to cover the fluidized bed cylinders and the blades.

Coating Over

Base Metal

Coating Over

Silica

FD 301930

Figure 7/7-29. Run 5-83-004 (100X) PWA 1447 Sputtered Onto PWA 658 Grooved Round

Bar. Grooves Were Filled With a Fused Silica Slurry

4. Silica Groove Filling Experiments

One of the machined fluidized bed test specimens was used in an experiment to determine

the effects of sputtering the cover material, PWA 1447, over the grooves filled with silica slurry.

The high-purity fused silica was mixed with a collodial silica binder to form a thick slurry, which

was then forced into the grooves machined about the OD of the test cylinder. The excess slurry

was removed by drawing a single-edge razor blade across the cylinder surface, then lightly

rubbing over the grooves with clean white paper. This final rub tended to smooth off the silica

deposit by transporting material from the high points to fill in any cracks or voids.

111-30
ORIGINAL PAGE IS

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FR-19009-1

Coating Over

Base Metal

Coating Over

Silica

FD 301931

Figure 111-30. Run 5-83-005 (100X) DC Triode — High Bias on Substrate. Substrate Is

Same Specimen Coated in Run 5-83-004. Surface was Left as Is

The specimen was heated at 170°F for 30 minutes to drive off moisture, then baked at 600°F

for one hour.

The deeper grooves proved relatively easy to fill by this method, but the shallow grooves

presented a problem. There seems to be very little adhesion of the silica slurry to the metal

cylinder and not much cohesion within the slurry itself. The primary mechanism which contains

the slurry in the grooves is probably friction between the silica particles and friction between the

particles and the side wall of the grooves. As shown in Figure 111-31, there is very little side wall

in the shallow grooves, hence it is difficult to contain the silica slurry in the shallow grooves.

Figure 111-32 illustrates the steps taken to fill the grooves with silica. The first attempt was

unsuccessful due to the aforementioned problems with shallow grooves. Extreme care was used

during the second attempt to avoid disturbing the fill material. After firing the cylinder, the fill

material (though still delicate) could be handled using moderate care.

111-31
Pratt & Whitney

FR-19009-1

Mag 10x

0.050 in. Wide

Deep Groove

0.035 in. Wide

Deep Groove

0.020 in. Wide

Deep Groove

0.050 in. Wide

Shallow Groove

0.035 in. Wide

Shallow Groove

Mag 1x

FD 302538

Figure III-31. Groove Geometry of Machined Fluidized Bed Test Cylinder

The above specimen was placed in the sputtering chamber and coated with approximately

10 mils of PWA 1447. The substrate was rotated at Vs rpm during coating to obtain a more

uniform deposit thickness about its circumference. This coating cycle was run using a triode

discharge. The substrate was biased with a negative voltage, which was switched periodically

from —10 vdc to -100 vdc in an attempt to deposit, as nearly as possible over the rough silica

surface, a "deposit free" coating.

Figure 111-33 shows the net result of this coating experiment. Apparently the stresses

induced by the developing coating overcame the holding force that the groove had on the silica

filler. Coating cracks and loss of filler cohesiveness were seen over every groove on the cylinder.

The result of this experiment showed that the silica slurry is not a good groove filler material.

111-32
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FR-19009-1

First Fill Attempt

Most of the Silica Was

Dislodged During Cleanup of Excess Slurry from

the Cylinder Surface

Second Fill Attempt

Extreme Care Was Taken

To Avoid Disturbing the

Slurry Deposit in Grooves

As Fired at 600° F/1 hr

After Firing the Cylinder May

Be Handled With Moderate Care

FD 302539

Figure 111-32. Appearance of Fluidized Bed Test Cylinder During Stages of Groove Filling

With a Silica Slurry

111-33
Pratt & Whitney

FR-19009-1

Run 5-83-005 Mag 1X

FD 302540

Figure HI-33. Run 5-83-005 — Fluidized Bed Test Specimen as Sputtered With

PWA 1447. The PWA 1447 Coating was Applied Over Grooves Filled With

Silica

5. Aluminum Groove Filling Experiments

The rest of the coating experiments were performed using aluminum as the groove filler

material. A method for filling the deep grooves in the cylinder with aluminum was developed and

used successfully to process three test specimens. A thin aluminum coat (approximately 0.2 mils)

was sputtered onto the grooved cylinder to form a well adhered, compliant layer on the grooved

wall. The steps illustrated in Figure 111-34 describe this process in detail.

The earlier difficulties encountered with the shallow grooves, during experiments with

silica, surfaced again with this method of groove filling. The lack of sufficient side-wall area on

the shallow grooves prevented swaging of the aluminum wire, which is necessary to anchor it in

place. None of the shallow grooves were successfully filled using aluminum wire. To fill these

shallow grooves, and also to clean up any gaps or machining defects in the wire-filled grooves, the

cylinders were sputtered with 6061 aluminum.

111-34
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Step A Step B Step C

Step D

Step A A Test Cylinder Which Has a Thin Coat of Sputtered

Aluminum Is Immersed in a Bath of Molten Wax. The

Wax Is Then Removed Flush to the Cylinder Surface

Using a Razor Blade and Fine Abrasive Paper

Step B The Cylinder Is Then Immersed In a Soln of NaOH

Which Removes All Aluminum Not Covered by Wax.

The Wax Is Then Removed by Immersion In Hot

Degreasing Solvent

Step C Aluminum Weld Wire (AMS 4043) Is "Peened" Into the

Grooves Using a Vibrating Electro Engraver With a

Rounded Tip. The Wire Diameter Is Approximately

0.005 in. Larger Than the Groove Width for Proper

Swaging Action

Step D Finally, the Excess Aluminum Wire Is Trimmed Back

Flush to Cylinder Surface on a Lathe and Finished

With Fine (No. 600) Abrasive Paper

FD 302541

Figure II1-34. Technique Used to Fill Grooves in Test Cylinders With Aluminum Wire

The first cylinder to go through this process was done in sputtering run 5-83-009. This

specimen was held by an uncooled, rotating fixture. The results are displayed in Figure 111-35.

Later temperature studies revealed that the substrate temperature reached approximately

1071 °F, apparently hot enough to cause the aluminum deposit to alloy with the substrate

material. A polished cross section of the coated cylinder wall revealed that some of the material

from the PWA 1447 substrate diffused outward into the coating as deposition was taking place.

This can be illustrated by the hardness data also shown in Figure 111-35. Normally, sputtered

111-35
Pratt & Whitney

FR-19009-1

aluminum is very soft, below 70 HV (Hardness, Vickers), as was shown in later tests. The

aluminum coating in this cross section was found to have a hardness of 543 HV, significantly

higher (harder) than that of the PWA 1447 base material. The alloy formed was very brittle, as

shown by violent chipping-off of the deposit encountered during later lathe machining of the

cylinder.

The remaining two cylinders were processed using a water cooled holder to reduce the

temperature of the substrate during deposition, thus reducing the possibility of alloying. Figure

111-36 depicts these two cylinders as they appeared after machining back to the original OD and

finishing with No. 600 SiC abrasive paper. Some striations can be seen in the aluminum, which

are caused by lathe cutting operations, but generally the finish looks entirely suited for

sputtering a "near-defect-free" overlay coating. Microscopic examination of the surface of both

cylinders revealed no cracks, voids, or other defects in the filler material, nor separations

between the filler and the groove walls.

An additional technique for filling the grooves was also investigated. Figure 111-37 shows

the product of a brief experiment to plasma-spray aluminum (AMS 4043) into the grooves of a

fluidized bed test specimen. This is the same cylinder which was previously used to experiment

with silica groove filling techniques and then coated with a cover of sputtered PWA 1447. The

coating was machined away and the remaining silica removed from the grooves. Several grooves

have spots which are filled with sputtered PWA 1447, in places where the silica filler fell out

prior to the sputtering of the cover layer.

The cylinder was machined back to flush with the original OD and finished with No. 600

SiC paper. The aluminum finish looks somewhat, but not much, rougher than the sputtered

aluminum filler. The leach-out holes in the grooves were not plugged prior to plasma spraying,

thus the holes are (for the most part) still very visible through the aluminum filler. It should be

noted that the actual spraying time required to fill the grooves was a matter of minutes, rather

than the hours required to fill them by sputtering. However, the sprayed deposit may be "gassy"

and perhaps may contain oxides or other contaminants.

An experiment was performed to determine the time required to leach out an aluminum

deposit from a passage (tunnel) of the same geometry that was used for the cooling passages of

the blade-tester blade. The size of these passages was expected to be 0.035 inch wide by

0.012 inch deep, with the length of the tunnel to be approximately one inch. A test fixture was

fabricated to simulate the cooling passage and a leach test was conducted.

The step-by-step procedure used in this investigation is illustrated in Figure 111-38.

Verification of full removal of the encapsulated aluminum was accomplished by comparing the

weight of the aluminum strip to that of the weight loss of the specimen after a period of time in a

7 molar NaOH solution. As a final check, when no further weight loss could be induced by

leaching, the nickel plate covering the aluminum strip (now a tunnel) was peeled back to expose

the area leached out (Figure 111-38). The test revealed that a time span of two hours would be

sufficient to remove all of the aluminum from passages with this geometry.

Prior to sputtering the PWA 1447 closeout layers on the two actual fluidized bed test

cylinders, a sputtering run was made to clean up the chamber after prior aluminum sputtering

cycles. The substrate selected to be coated with PWA 1447 in this run was a cast PWA 1447 rig

bar. The coating cycle was run using the same parameters — triode mode, negatively biased

substrate. The run was continued for two hours and fifteen minutes and produced a coating

approximately 0.003 inch thick.

111-36
Pratt & Whitney

FR-19009-1

ORIGINAL

Of POOR QUALITY

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111-37

4029C
ORIGINAL PAGE IS

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Pratt & Whitney

FR-19009-1

Mag 1X

Run 5-83-016 Run 5-83-017

Typical Appearance of Aluminum Filled Grooves

After Lathe Machining Back to the Cylinder

Surface and Finishing With No. 600 Grit

Abrasive Paper

Note: These Two Cylinders Were Coated While

Held on a Water-Cooled Copper Fixture

Mag 4x

Figure HI-36. Sputtered Aluminum in Fluidized Bed Test Cylinder Grooves

FD 302543

A cross section was cut from the middle of this bar and processed for metallurgical

examination. Microprobe scans of these areas showed that the sputtered coating was similar in

chemical composition (within the tolerance range of the microprobe) to the PWA specifications

for this material, as depicted in Table III-5. Metallurgical evaluation also showed that, although

this was a cleanup run for the chamber with a newly reinstalled target and some contamination

was to be expected, the interface and the sputtered deposit looked clean and virtually defect-free.

The few spots found along the interface were actually bits of embedded silicon carbide, remnants

of processing the rig bar by grit blasting at some earlier date.

111-38
Pratt & Whitney

FR-19009-1

Mag 1X

The Cylinder Was Machined Back Flush to the Original

OD Surface.

FD 302544

Figure 7/7-37. Plasma Sprayed Aluminum (AMS 4043) in Grooves on PWA 1447 Test

Cylinder

Table III-6 contains the hardness values for the same cross section. A Zwick microhardness

tester was used to obtain hardness values of the coating, to further illustrate its similarity to (or

difference from) material known to be PWA 1447. The data presented here indicates that the

hardness of the sputtered deposit does not differ significantly from that of the cast PWA 1447.

6. Sputtering of PWA 1447 Cover Material

The fluidized bed test cylinder which had grooves filled with plasma sprayed aluminum was

the first specimen selected for application of the PWA 1447 close-out (cover) layer. It was

intended to use this specimen as a trial run to evaluate the behavior of the sputtering system (i.e.,

check out control parameter settings) before committing one of the new cylinders. A

thermocouple was wired to the OD of this cylinder and the surface temperature was monitored

during coating deposition. Various sputtering parameters were adjusted to formulate a cycle

which would deposit a defect-free, well bonded coating in a reasonable amount of time, while

minimizing the process temperature in order to limit interdiffusion of the PWA 1447 and the

aluminum.

111-39
Pratt & Whitney

FR-19009-1

Two Holes, 1 in. Apart, Were Drilled Into a Stainless Steel

Baseplate. One Hole Had a Diameter of 0.035 in., the

Other 0.014 in.

A Strip of Aluminum, 0.035 in. Wide, 0.012 in. High, and 1.10 in.

Long Was Epoxied Onto the Baseplate So as to Span the

Distance Between the Two Holes

The Fixture Was Then Nickel Plated. A Piece of Tape Covered

the Backside of the Baseplate To Keep the Leaching

Holes Open

The Assembly Then Was Immersed in a NaOH Solution, 7 Molar

Concentration, and the Time Required to Completely Leach Out

the Aluminum Through the Holes Was Charted

""It Was Determined That All of the Aluminum Would Be

Leached Out in 2 hr

FD 302545

Figure 111-38. Aluminum Leach Rate Study

This cylinder was held in the same water-cooled cylindrical copper fixture used to deposit

aluminum earlier in the program. A twisted double strand of 18 AWG (0.040 inch diameter) bare

copper wire was wrapped around the body of the holder to provide intimate contact with the ID

of the test cylinder for more efficient heat transfer.

111-40
Pratt & Whitney

FR-19009-1

Table 7/7-5. Microprobe Results of Sputtering Experiment 5-83-018 — Evaluation of

Coating and Substrate Chemistry

Element

Ni

Cr

Co

Mo

Al

Ta

W

Hf

Ti

B

Zr

PWA Spec

Bal

8.4

10

0.65

5.5

3.0

10

1.4

1.0

0.15

0.05

Substrate

Bal

10.3

9.4

0.7

6.0

2.3

9.2

0.9

1.0

_

—

Coating

Bal

12.1

10

0.7

4.7

2.8

8.6

0.8

1.0

—

—

Table 777-6. Microhardness Testing of Sputtered PWA 1447 Run 5-83-018

'Hardness Readings, in Hv, as Determined by a Zwich Microhardness Tester Using a

Vicker's Pyramid Indenter Under a 300 gm Load.

•Values Shown are the Averages of 10 Readings at Each Location.

Cast PWA 1447 Base 395 HV (average pyramid height of 0.037 mm)

Sputtered PWA 1447 429 HV (average pyramid height of 0.036 mm)

An average thickness of 4 mils (0.004 inch) was obtained during 4.5 hours of deposition.

During the deposition cycle, inordinate pressure fluctuations were experienced, caused by the

outgassing of heated surfaces within the coating chamber. Upon removing the cylinder from the

sputtering system, it was discovered that much of the coating over the grooves contained nodular

defects. It is believed, that these defects were caused by the release of gases trapped in the

aluminum filler during the plasma spray application. Later coating runs using cylinders with

grooves filled with sputtered aluminum and aluminum wire showed considerably fewer defects of

this type. It should be noted that due to limitations imposed by the attachment of the

thermocouple, the cylinder was not rotated. Also, the metal sheath of the thermocouple placed

the cylinder at ground potential which precluded the use of biasing to improve the deposit

uniformity. Figure 111-39 shows the condition of this cylinder as coated.

It was decided to leach out the aluminum at this point so that the remainder of the 10 mil

(0.010 inch) thick coating could be applied under less restrictive conditions. The cylinder was

immersed in a 7 molar NaOH solution maintained at 140 to 160°F. Prior experiments had

determined that cooling channels of geometry similar to those in this cylinder would require

approximately two hours to leach free of aluminum using this solution. However, this cylinder

had seen prior service as a substrate in a coating experiment where PWA 1447 was sputtered over

silica-filled grooves. The silica had fallen out in many places prior to and during the sputter

coating cycle. The voids left by the wayward silica were filled by the PWA 1447 material which

ultimately blocked the channel in several places. This blockage limited the flow, hence the

effectiveness, of the NaOH leaching solution. Thus, rather than the two hours leaching time

indicated by past experiments, this specimen required almost 24 hours of soaking along with

frequent ultrasonic cleaning cycles (this problem would not be encountered in a new cylinder).

After this time period, some small bubbles of gas were still being generated by the chemical

reaction; but for the purposes of this initial experiment, it was determined that enough of the

aluminum had been removed to continue processing of the cylinder. The leached out cylinder is

shown in Figure 111-40.

111-41
Pratt & Whitney

FR-19009-1

As Coated 1X

FD 267902

Figure 7/7-39. Plasma Sprayed Aluminum Fluidized Bed Test Cylinder as Coated With

PWA 1447 in Run 5-83-019

Coating of the cover layer on this cylinder was continued in sputter run 5-83-020. The

specimen was fixtured on the same water-cooled holder as before, but this time no copper wire

was added to fill the space between the OD of the holder body and the ID of the cylinder. The

cylinder was inverted with respect to the vertical orientation observed during its first coating in

order to obtain a more uniform coating thickness throughout its entire length. The thermocouple

was removed which permitted a bias voltage of —25 vdc to be placed on the substrate. Also, for

this second layer the target voltage was increased to 1.5 kvdc, which was 1.5 times the target

voltage employed previously. The coating cycle was continued for five hours. The coating

deposited during this experiment contained many nodular defects over the grooves where

incomplete removal of the filler material probably resulted in the vaporization of the over-heated

aluminum. Figure 111-41 illustrates the appearance of the as-coated cylinder from run 5-83-020.

The next cylinder to be coated with a cover of PWA 1447 was one of the fluidized bed test

cylinders which had the shallow grooves filled with sputtered aluminum and the deep grooves

filled with aluminum wire. The parameters used for this coating run were adjusted to minimize

the temperatures generated by the sputtering process to avoid alloying the aluminum. Despite

precautions taken to reduce the heat, one side of the cylinder developed several nodular defects

along the grooves. Also, during subsequent processing to leach the aluminum from the grooves,

much of the PWA 1447 coating debonded from the cylinder, with the remaining coating easily

removed by fingernailing. On this side of the cylinder even the "nodule free" deposit over the

aluminum-filled grooves looked wavy, as if it had been deposited over a flowing surface. Figure

111-42 illustrates the as-coated condition of this cylinder.

111-42
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As Leached IX As Leached 4X

FD 267755

Figure 111-40. Plasma Sprayed Aluminum Fluidized Bed Test Cylinder After Leaching Out

Aluminum Filter from Grooves

To eliminate the localized heating which was believed to have caused the nodular defects, a

modification was made to the sputtering chamber to permit the rotation of the cylinder about its

axis during coating, while maintaining the option to water-cool the substrate fixture. This was

accomplished by using a magnetically sealed rotatable vacuum feedthrough which sealed and

rotated the 0.5 inch diameter stainless steel support tube connected to the substrate holder. The

cooling water was fed to this tube using a fixed coupling fitted with a well greased 0 ring. This

arrangement allowed the water feed and drain lines to remain stationary while supplying running

water to the revolving substrate assembly.

The next cylinder selected for coating was the second fluidized bed test cylinder with

sputtered and wire-formed aluminum within its grooves. This specimen was mounted on the

modified water-cooled rotating holder. The double stranded copper wire was again used to

provide intimate contact between the substrate and cooled holder surfaces. Again, every attempt

was made to limit the temperature encountered by the substrate. In addition to temperature

limiting procedures employed in prior experiments, this run omitted the typical predeposition

sputter-clean (reverse sputtering) cycle used to outgas and ion scrub the substrate and target

surfaces. To compensate for the omitted procedures, the chamber was pumped for three days

under moderate heat (<400°F) to remove as much gas as possible from the surfaces within.

111-43
Pratt & Whitney

FR-19009-1

5-83-020 Mag IX

Second Coating of PWA 1447 Cover

FD 267756

Figure 111-41. Appearance of Fluidized Bed Test Cylinder After Second Coating of

Sputtered PWA 1447

The coating from this run looked good except for a small area at the bottom end of the

cylinder which contained a few nodular defects. During removal of the cylinder from the holder,

it was noted that the copper wire used for heat transfer had not reached the cylinder bottom. One

of the aluminum plugs used to fill the holes in the groove bottoms prior to sputtering of the

aluminum filler protruded past the cylinder ID wall and blocked the path of the wire as it was

wound in from the top of the fixture. This severely limited the heat flow away from the bottom

third of the cylinder during coating. The appearance of this cylinder as-coated is shown in Figure

111-43.

The second coating on this cylinder was applied with the aluminum groove filler still in

place. Prior to fixturing, the cylinder was put on a lathe and No. 600 grit SiC paper was used to

clean up the OD to provide a smooth surface for the final coat. Also, the ID was treated with

No. 120 grit SiC paper to remove all protruding aluminum plugs. For this run, a single strand of

copper wire was wound around the substrate holder and forced against the cylinder wall for

maximum contact. The cylinder was inverted to provide more uniform coating thickness. During

the coating cycle the water-cooled substrate holder was rotated at 1 rpm.

111-44
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Best Side

Sputtering run 5-83-021

Mag IX Worst Side Mag IX

PWA 1447 sputtered over a PWA 1447 fluidized bed test cylinder

having cooling grooves filled with aluminum wire and sputtered

aluminum.

FD 267757

Figure 111-42. Two Sides of the Fluidized Bed Test Cylinder Coated in Sputter

Run 5-83-021

The resultant coating from this experiment looked good and was virtually defect-free.

However, much of the coating flaked off later during the aluminum leaching process, indicating

poor bonding. Figure 111-44 shows the cylinder before and after the final coating in run 5-83-023.

Figure 111-45 illustrates the cylinder condition after leaching.

The underside of one of the debonded coating strips from the cylinder coated in run

5-83-023 was analyzed by X-Ray Energy Spectroscopy (XES) to determine the chemical makeup

of the coating at the substrate interface. The results, as displayed in Figure 111-46, did not show

abnormal indications which might lead to coating debonding. Analyses of the surface of the

cylinder itself in an area under a debonded strip showed no significant variation from

specifications for PWA 1447. A deposit of unleached aluminum near one of the cooling holes in

the cylinder was also charted by XES. This chart, displayed in Figure 111-47, shows a significant

amount of copper which could only have come from the copper wire used for heat transfer.

Apparently, the aluminum plugs extending through the cylinder wall to the ID contacted the

copper wire and alloying of the two occurred even at the relatively low temperatures seen close to

the water-cooled substrate wall. The copper alloyed with the aluminum, along with traces of

111-45
Pratt & Whitney

FR-19009-1

other elements from the PWA 1447 base material, caused a drastic slowdown of the leaching

process. An analysis of this alloyed aluminum is presented in Table III-7 along with a similar

analysis of the aluminum wire used to plug the cooling holes.

Run 5-83-022 As-Coated Mag IX

FD 267903

Figure 111-43. PWA 1447 Sputtered Over PWA 1447 Test Cylinder, Run 5-83-022

(First Coat)

7. Bond Strength Investigation

To understand the lack of bonding strength between the PWA 1447 sputtered coating and

substrate, an effort was initiated in the Materials Laboratory to evaluate this problem.

Acting on the suggestion from NASA that the thermal expansion difference between the

aluminum and the PWA 1447 may have contributed to the coating debonding noted during

leaching, an experiment was designed to duplicate these conditions in a controlled environment.

A 0.062 inch hole was drilled approximately 0.25 inch deep into a PWA 1447 rig bar section,

which measured approximately 0.5 inch diameter by 0.75 inch in length. Aluminum wire (4043),

0.060 inch diameter, was then forced into the hole by vibro-peening. The aluminum was ground

down flush with the surface of the host bar section and the composite face was finished to

No. 320 grit SiC paper to produce a smooth, flat surface. Next, a 0.01 inch thick layer of nickel

was plated over the specimen. The plated specimen was immersed in boiling water for one hour

to effect a "worst case" simulation of the heat derived from the aluminum leaching process. After

this exercise, the sample was examined for any bulge (or depression) in the nickel plate over the

aluminum wire, which might have been formed by thermal expansion mismatching between the

PWA 1447 and the aluminum. No abnormalities were noted in the nickel plate. As a further

check, the specimen was placed in an oven at 200°C. After 30 minutes the nickel plate still

appeared flat and featureless. A cross section of the specimen was prepared and microscopically

examined. Still no distinct bulge could be seen, but, an area approximately one-half the size of

111-46
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FR-19009-1

the total flat surface area and centered over the aluminum wire had debonded and remained

slightly lifted from the PWA 1447 surface. The progressive steps involved in this experiment are

illustrated in Figure 111-48.

Run 5-83-023 As Fixtured IX Run 5-83-023 As-Coated Mag IX

FD 267758

Figure 111-44. PWA 1447 Sputtered Over PWA 1447 Test Cylinder, Run 5-83-023

(Final Coat)

In the next experiment a 0.035 inch hole was drilled all the way through a 0.25 inch thick

section of a PWA 1447 rig bar. A piece of aluminum wire was thinned down to 0.35 inch using a

solution of NaOH then drawn through the hole filling it completely. The two faces of the bar

section were then polished to No. 600 SiC paper to obtain smooth, flat surfaces. The specimen

was then coated with 0.001 inch of nickel in a RF Diode sputtering system. No surface

irregularities were noted after coating. One face of the specimen was ground down past the nickel

coating to expose the PWA 1447/aluminum composite surface. The aluminum was leached out

through this hole using a 7 molar NaOH solution held at 140 to 160°F. The leaching process

required two hours to complete because of the tendency of the reaction to slow drastically as the

excavated hole deepened. This probably occurred because the gas generated by the chemical

reaction formed a pocket within the hole and restricted the passage of fresh solution to the

aluminum surface. It should be noted that the orientation of the specimen to be leached is of

great importance in the leaching process. If the cylinder or blade is placed within the NaOH

solution in such a manner that gas bubbles formed must be forced down through a channel, the

natural buoyancy of the gas could tend to hold the bubbles inside the inverted hole and slow or

even stop the chemical reaction. Any complex structure probably would contain several channels

having an "inverted" orientation, so some form of rotation of the specimen during the leaching

process would be required.

111-47
Pratt & Whitney

FR-19009-1

Run 5-83-203 Mag IX

Shortly After Start of Leaching Cycle

FD 267759

Figure 111-45. Appearance of Fluidized Bed Test Cylinder Coated in Run 5-83-023, After

NaOH Leach

The steps involved in this experiment as well as the results are illustrated in Figure 111-49.

This time a bulge did appear in the coating over the hole as predicted. This specimen was next

prepared for metallurgical examination of its cross section. Microscopic viewing of the polished

cross section, as shown in Figure 111-50, revealed that the coating had separated in the area only

over the aluminum-filled hole. A gap between the coating and the walls of the hole measured

0.0006 inch wide.

A series of experiments were also run to determine other possible causes for poor coating

adherence and to develop a means to overcome this problem. In these experiments, films of

PWA 1447 and Ni-200 (commercially pure) were sputtered onto strips of PWA 1447 and Ni-200.

Adherence testing of RF diode sputtered Ni-200 on PWA 1447 and Ni-200 strips showed no

debonding of the coating up to 10,000 psi, the limit of the tensile tester used. Adherence testing of

DC triode sputtered PWA 1447 coating on PWA 1447 and Ni-200 strips also showed no

debonding to 10,000 psi. The only coating that showed any sign of failure at these limits was the

Ni-200 which was RF diode sputtered onto a vapor blasted PWA 1447 surface. Figure 111-51

shows a typical coated strip, in this case PWA 1447 sputtered onto a PWA 1447 substrate which

was first vapor blasted, as it emerged from the tensile tester. The bonded pin was still attached

indicating no debonding occurred within the limits of the test equipment.

111-48
Pratt & Whitney

FR-19009-1

Counts

Coatings, 6/22/83

75 sec

2000

1600

1200

800

400

Coating Interface

W

AL

Nl

KEV

5

XES

10

FD 267901

Figure 111-46. Results of XES Analysis for Underside of the Debonded Coating

Four strips, each approximately 1.5 inches long by 0.25 inch wide by 0.04 inch thick were

bend tested after coating with DC triode sputtered PWA 1447. Two of the strips were made of

PWA 1447 material and two of Ni-200. One each of the two materials was prepared by vapor

blasting the surface and the other by polishing down to No. 600 grit SiC paper. The strips were

bent 90 degrees over a 0.28 inch diameter mandrel. In each case the PWA 1447 coating adhered

well to the nickel strip and spalled from the PWA 1447 strip (severely spalled from the polished

PWA 1447 strip). Figure 111-52 illustrates the appearance of these strips prior to bending. Figure

111-53 shows the results of the bend testing.

As a final attempt to resolve the bonding problem between the PWA 1447 substrate and

cover layer, a nominal 0.001 inch thick coating of nickel was deposited between the two layers.

Specimen appearance and surface topography are shown in Figure 111-54. The nodular

topography in each case is typical of a cold substrate. The nodular growth is more pronounced

over the aluminum wire filler which had a rougher surface than the PWA 1447 land. The nickel

closeout layer was essentially nonadherent, as shown in Figure IH-55.

111-49
Pratt & Whitney

FR-19009-1

Counts

Coatings, 6/22/83

56 sec

2000

1600

1200

800

400

Unleached AL Near Hole in Groove

AL

CU

SI - W CR

Tl

KEV

5

XES

10

FD 267760

Figure 111-47. Results of XES Analysis for Unleached Aluminum

Due to the difficulties in depositing a well adherent cover layer on the PWA 1447 fluidized

bed test cylinders, it was determined that the technology required to perform this task was not

available at this time. An expensive development program would be required to achieve the task.

Such a program was out of the scope of the advanced turbine study. A recommendation was made

to NASA to cease work on the advanced convectively cooled blade and redirect the program to

evaluate the hot core blade.

Appendix A contains a summary of the sputtering work conducted as part of the ATS

program. An overview of the state of the art of the sputtering technology is presented, as well as

the recommended development program which would be required to resolve the lack of adhesion

between the PWA 1447 substrate and the cover layer.

D. TASK 2 — SUBTASK 203, SPECIMEN TEST

This subtask consisted of the fluidized bed testing of the test cylinders simulating the

advanced convectively cooled turbine blade.

111-50
Pratt & Whitney

FR-19009-1

Table 7/7-7. Comparison of Aluminum Left in Cooling Channel and 4030 Aluminum

Wire

Element

and Line

AL KA

SI KA

TI KA

CR KA

CO KA

NI KA

CU KA

W KA

Weight %

33.30

5.16

0.68

5.95

2.60

12.40

26.97

12.93

Atomic %

53.74

7.99

0.62

4.99

1.92

9.20

18.48

3.06

Precision

2 Sigma

0.91

0.34

0.12

0.41

0.34

0.81

1.42

0.85

K-Ratio

0.1421

0.0211

0.0062

0.0578

0.0271

0.1277

0.2566

0.0661

ITER

5

TOTAL 100.00

NORMALIZATION FACTOR: 0.705

XES Analysis of Material Left in Channel After NaOH Leaching

STANDARDLESS EDS ANALYSIS

(ZAF) CORRECTIONS VIA MAGIC V)

Element

and Line Weight % Atomic %

Precision

2 Sigma K-Ratio

AL KA

SI KA

97.35

2.65

97.46

2.54

1.35

0.40

0.9904

0.00%

TOTAL 100.00

NORMALIZATION FACTOR: 0.981

XES Analysis of 4030 Aluminum Wire Used to Fill Cooling Channels

4041C

By exposing the cylinders (which were fabricated with several coolant groove-land geometry

combinations as shown in Figure III-14) to very rapid temperature changes in the fluidized bed,

the high thermal gradients produced by the blade tester rig and the SSME were to be

approximated. After this thermal exposure, posttest evaluation of the hardware should have

allowed the selection of the optimum coolant groove-land geometry combination.

The rapid temperature changes were to be achieved by rapidly submerging the test

specimens in alternate hot and cold silica sand baths. The temperatures in these baths were to

have been closely monitored and controlled to ensure the proper AT between both baths.

Due to the inability to fabricate the test cylinders, the fluidized bed tests were not

performed.

E. TASK 3 — SUBTASK 301, AIRFOIL DESIGN AND ANALYSIS

The design of the advanced convectively cooled blade was conducted in parallel to the

design and fabrication of the fluidized bed test cylinder to meet the hardware delivery dates

required by the program.

111-51
ORIGINAL PAGE IS

OF POOR QUALITY

Pratt & Whitney

FR-19009-1

Step 1. A 0.062 inch diameter hole was drilled

into a bar section of PWA 1447. The 0.25

inch deep hole was then filled with 4043

aluminum wire. The face was polished

smooth and flat.

Step 2. The specimen was coated with 10 mils

(0.01 inch) of plated nickel.

Step 3. The specimen was immersed in boiling

water for one hour.

Step 4. Finally, the specimen was heated for 30

minutes in an oven at 200°C.

No bulge over the aluminum wire was noted during

or after any of the above procedures.

FD 267762

Figure 111-48. First Experiment to Investigate Effects of Thermal Expansion Mismatch

Between Aluminum and PWA 1447

111-52
ORIGINAL PAGE SS

OF POOR QUALITY

Pratt & Whitney

FR-19009-1

A 0.035 inch hole was drilled completely

through a 0.25 in. thick bar section of PWA

1447. A 0.035 inch diameter aluminum wire was

drawn through the hole and the composite face

was polished flat.

2. The specimen was next coated with 1 mil

(0.001 inch) of sputtered nickel in an RF diode

discharge.

3. After two hours in a NaOH aluminum leaching

solution maintained at 140 to 160°F a definite

bulge can be seen in the nickel cover over the

aluminum wire.

FD 267763

Figure 111-49. Effect of Thermal Expansion Mismatch Between Aluminum and PWA 1447

A number of studies were performed in support of this effort. One of the first studies

evaluated the effect of coolant film coefficient on maximum metal temperature for typical rig

conditions and coolant geometry. The results of this study are shown in Figure 111-56. This study

showed that increasing the coolant film coefficient above 4000 Btu/hr ft2 °F has a small payoff in

terms of reduced metal temperature.

111-53
Pratt & Whitney

FR-19009-1

Plated Ni

Sputtered Ni

I Separation

Sputtering Run 6-83-083 MN 30806-1 Mag 200X

The coating separation at the walls of the hole was measured

optically by microscope. A gap of 0.0006 inch was recorded at

each side.

FD 267764

Figure IH-50. Cross Section of Bulged Area Over Aluminum Wire Inlaid in PWA 1447

Bar

The pressure distribution and gas bending stresses from the aerodynamic analysis of the

SSME blade in the rig are shown in Figure 111-57. The low level of gas bending stress confirms

that the major component of stress is due to thermals. Our blade design has even lower bending

stresses because of the higher moments of inertia associated with a cooled airfoil cross section.

The pressure distribution was used for calculating preliminary heat transfer film coefficients for

the rig environment.

Figure 111-58 shows the preliminary airfoil cross section selected for detailed evaluation.

The blade was designed for ease of manufacturing (constant airfoil section without twist or taper,

and relatively large leading and trailing edge diameters). These simplifications were made in

order to meet cost and deadline commitments. Refinements in the aerodynamics, such as varying

the airfoil cross section from root to tip and reducing the trailing edge thickness, are required in a

production blade design. This can be achieved in a cost effective manner when production tooling

is used.

111-54
Pratt & Whitney

FR-19009-1

Adherence Testing of Sputtered PWA 1447 on PWA

1447 Strip Mag 1X

Polished Vapor Blasted

Adherence Test Results, Sputtered PWA 1447 on PWA

1447 Strip With One Side Vapor Blasted and the Other

Side Polished.

FD 267765

Figure 7/7-57. Adherence Testing of Sputtered Coatings on Test Strips

An example of the nodal heat transfer models used in the analysis of the rig heat transfer,

for both the methane cooled and hydrogen cooled parts, is shown in Figure 111-59. The model

depicts one of the coolant passages in the airfoil wall. Since all the passages are the same, only

one is modeled. The maximum/average metal temperature for the hydrogen cooled foil is

1359°F/550°F and the maximum thermal strain is 0.40%. The methane cooled foil has a

maximum/average metal temperature of 1430°F/966°F and a maximum strain range of 0.225%.

These are for a rig hot gas temperature of 1700°F and coolant of 70°F. A quantity of coolant

equal to 2% of mainstream flow was used.

111-55
Pratt & Whitney

FR-19009-1

#5 PWA 1447 Strip, Vapor Blasted

#4 Ni-200 Strip, Vapor Blasted

#3 PWA 1447 Strip, Polished

#2 Ni-200 Strip, Polished

#1 PWA 1447 Strip, 1/2 VB, 1/2 Pol

FD 267766

Figure 7/7-52. Appearance of Test Strips After Sputtering With 0.01 Inch Coating of

PWA 1447

After several iterations the rig blade final outside contour was defined. The final blade

eontour and the resultant mainstream pressure distribution are shown in Figures 111-60 and

111-61, respectively. The leading edge and trailing edge diameters are 0.15 and 0.07 inch,

respectively. The contour was designed with the intent of providing an easily manufacturable

blade with no mainstream separation when running in the rig.

111-56
ORIGINAL PAGE IS

OF POOR QUALITY Pratt & Whitney

FR-19009-1

Mag 4X Mag 500X

Tab #2 PWA 1447 on Polished Ni-200

Mag 4X Mag 500X

Tab #3 PWA 1447 on Polished PWA 1447

IHH

Mag 4X Mag 500X Mag 4X Mag 500X

Tab #4 PWA 1447 on Vapor Blasted Ni-200 Tab #5 PWA 1447 on Vapor Blasted PWA 1447

FD 267767

Figure 7/7-53. Results of 90 Degree Bend Testing of Coated Test Strips

The aerodynamic losses associated with trailing edge blockage are an important consideration when evaluating the overall benefit of operating cooled blades at a higher turbine inlet

temperature. Since the rig blade contour and coolant passages will be machined from a bar, a

relatively large trailing edge was used for ease of manufacture, but production hardware blades

will be cast with integral slots. With precision castings smaller trailing edge diameters will be

possible, as shown in Figure 111-62.

111-57
Pratt & Whitney

FR-19009-1

ORIGINAL PAGE JS

OF POOR QUALITY

Ni-200 Coating Over Land Surface Ni-200 Coating Over Aluminum Surface

fD 267792

Figure 111-54. Specimen from Sputtering Run 6-83-086 SEM Photographs 500X

Magnification

111-58
Pratt & Whitney

FR-19009-1

As Coated With Ni-200

in Sputter Run 6-83-086

Deep Grooves Had Been

Filled With Aluminum

Wire Prior To Coating

fii i ii § n •

Appearance of Specimen

After Leaching of Aluminum

Filler Material With a

NaOH Solution

Appearance of Specimen

After Attempts To Section

for Metallurgical Review

FD 267793

Figure 7/7-55. Adherence Failure of 0.001 Inch Thick Nickel Closeout Layer

In order to determine the size of the central core that runs up the center of the blade, a

study was undertaken to determine the effect of wall thickness on thermal strains. The results of

the detailed thermal and stress analysis are shown in Figure 111-63. The study showed that the

thinner the wall, the lower the thermal strains. The minimum wall thickness was chosen to be

0.055 inch which resulted from the structural requirement to limit the stresses and deflections

caused by the coolant-to-mainstream pressure differential.

111-59
Pratt & Whitney

FR-19009-1

r

0.010 in.

Hot Gas Max Metal Temperature Location

Coolant Coolant

I——0.045 in. -I Slot Depth

Variable

T = 1700 °F (H + OJ

hot gas 2 T

H = 10,500 Btu/hr ft2 °F hot gas '

2% of Mainstream for Coolant

20001—

1500

Temperature - °F 1000

500

hot ga>

T Metal Max Local

T Metal Hot Surface Ave

f f

0.023 in. Deep 0.0115 in. Deep

Slot Slot

4000 8000 12,000 16,000

H Coolant Film Coefficient - Btu/hr ft* °F

20,000

FDA 301934

Figure 111-56. Effect of Coolant Fluid Coefficient on Blade Metal Temperature for Rig

Conditions

HI-60
1.00

0.95 —

Pratt & Whitney

FR-19009-1

SSME 1st Blade Midspan in NASA Rig

X(CAS)/BX

(Function of axial chord length)

Gas Bending Stress

Maximum Bending Stress

on Leading Edge

40 50 60

Percent Span

70 80 90 .100.

FD 253260

Figure 111-57. External Pressure Distribution

1. Preliminary Life Analysis of ATS Cooled Rig Foils

The methane cooled foil and a hydrogen cooled foil were designed and evaluated for testing

in the high-pressure NASA cyclic, nonrotating (blade tester) rig. The NASA rig produces a

severe cyclic environment as evidenced by the very high acceleration and deceleration rates and

hot gas film coefficients of 6,000 to 15,000 Btu/hr ft2 °F. Current space shuttle uncooled

HI-61
Pratt & Whitney

FR-19009-1

directionally solidified uncoated (DS) Mar-M 246 blades show severe cracking after 5 to 25 cycles

in this rig environment.

1.0 I—

0.9

0.8

0.7

0.6

Y/BX

0.5

0.4

0.3

0.2

0.1

0.0

-0.1

— Chord Angle = 19.5963

I I I I I I I I I I I I

0.0 0.1 0.2 0.3 0.4 0.5 0.6 0.7

X/BX

0.8 0.9 1.0

FDA 301935

Figure 7/7-58. Advanced Turbine Study Preliminary Blade Cross Section

The cooling scheme used (Figure 111-64) is a series of chordwise coolant passages which are

fed from a radial core through small leading edge holes. The passages discharge at the trailing

edge. The coolant passages are 0.0115 inch deep and 0.035 inch wide separated by 0.045 inch

lands. The passage lies 0.010 inch from the outside surface, thus minimizing the large metal

temperature drop through the wall in this high heat flux application. The foils flow the

equivalent of 2% of total mainstream flow.

111-62
257

265

256

275

276

Sputtered Cover

249 243

Pratt & Whitney

FR-19009-1

283 289

24

28

29

30

31

32

45

46

47

48

49

50

51

52

53

54

55

56

57

23

27

33

34

35

36

58

59

60

61

62

63

64

65

66

67

68

69

70

22

26

37

38

39

40

71

72

73

74

75

76

77

78

79

80

81

82

83

21

25

41

42

43

44

84

85

86

87

88

89

90

91

92

93

94

95

96

14

20

-323

-320

97

98

99

100

101

102

103

104

105

106

107

108

109

13

19

12

18

11

17

10

16

Coolant

Passage

110

111

112

113

114

115

116

117

118

119

120

121

122

123

124

125

126

127

128

129

130

131

132

133

134

135

136

137

138

139

140

141

142

143

144

145

146

147

148

149

160

151

152

153

154

155

156

157

158

159

160

161

9

15

310-

313-

162

163

164

165

166 .

167

168

169

170

171

172

173

174

4

8

242

241

240

239

175

176

177

178

179

180

181

182

183

184

185

186

187

3

7

238

237

236

235

188

189

190

191

192

193

194

195

196

197

198

199

200

2

6

234

233

232

231

201

202

203

204

205

206

207

208

209

210

211

212

213

1

5

230

229

228

227

214

215

216

217

218

219

220

221

222

223

224

225

226

308

0.

300

290

FDA 301936

Figure 111-59. Nodal Breakup of Wall With Coolant. Passage

The cyclic life goal for the blades is 100 cycles. The preliminary life analysis shows the

methane cooled foil with a life of 200 cycles and the hydrogen cooled foil with a life of 560 cycles

(using average material properties). The coolant is designed for positive outflow in order to

provide a graceful wearout mode.

111-63
Pratt & Whitney

FR-19009-1

i

0)

T3

a

CO

u.

5.

x

i

1

<

CO

o 8

<q

d

CVJ

d

p

d

<p

d

evi

d

p

d

«l

S

t•*j

co

"8

m

§

in

o

in

o

o

p

in

CD

3 |

c\i in

I

s s

p in

i-' in

CM

CM

(O

CM

in

d

CVI

o

CM

CM

CM

s 8P

d

•o

^

n n

§

o

111 s

13

BETA1

S

DBETA1

CM

oc

O 111

II 1 (U

BETA2

S

DBETA2

0

B2GAGE

GAMM

A

X

00

UJ

N

111-64
Pratt & Whitney

FR-19009-1

ATS First Blade Mean Section

T882 Pressure Distribution - P/PT vs X/BX

Inlet Mach Number = 0.0

Inlet Gas Angle = 61.700 >

HLE/H1 = 1.000

HTE/HLE = 1.010

H2/HTE = 1.000

H2/H1 = 1.010

0.178

61.700

Exit Mach No.

Exit Gas Angle

BX

Loss (DPT/PT)

Turning

Incidence

0.396

25.500

0.738

0.020

92.800

3.300

0.396

25.500

0.020

P/PT

1.01—

0.9

0.8

0.7

0.6

0.5

0.4

0.3

-0.20 0.00 0.20 0.40

X/BX

Axial Chord

0.60 0.80 1.00

FDA 301938

. Figure 111-61. Advanced Turbine Study Rig Blade Pressure Distribution

HI-65
Pratt & Whitney

FR-19009-1

0.07 in.

Trailing

Edge

Cover

In Line Coolant

Passages

(Machined)

nt

i

I

0.045 in.

Coolant Passage Width

Locally Reduced

at Trailing Edge

Radially Staggered

Coolant Passages

(Cast In)

Proposed Rig

Trailing Edge

Possible Production

Trailing Edge

Enlarged Sections A-A Shown Above

FDA 301939

Figure 7/7-62. Method for Reducing Trailing Edge Diameter in Production Hardware

Figures 111-65 and IH-66 show the transient hot gas temperature and pressure that the

NASA rig generates. The steepest parts of the acceleration and deceleration are equal to

9000°F/second, which is significantly more severe than a typical jet engine transient. Table III-8

lists the hot gas coolant heat transfer film coefficients and temperatures at various locations

around the foil. Note that the film coefficients are significantly higher than those encountered in

a jet engine environment. The high film coefficients combined with the rapid transient causes a

severe thermal shock for the foils. This translates into high thermal strains in the foil wall.

Figures 111-67 through 111-71 show this strain versus metal temperature for five locations around

the hydrogen cooled airfoil. The highest strain range of 0.684% occurs on the suction side at 86%

111-66
Pratt & Whitney

FR-19009-1

axial chord just aft of the rear core cavity. This corresponds to a cyclic life of 560 cycles. Figures

111-72 through 111-76 show the strain versus metal temperature for five locations around the

methane cooled foil. The highest strain range of 0.82% occurs on the suction side at 86% axial

chord just aft of the rear cavity. This corresponds to a cyclic life of 200 cycles. Figure 111-77

summarizes the strain range and cyclic life for the hydrogen cooled and methane cooled foils.

In general, the thermal strains in the hydrogen cooled foil should be higher than the

methane cooled foil because of the higher heat flux in the hydrogen cooled foil. But in the thick

wall section aft of the rear core cavity, the methane configuration shows higher strains than the

hydrogen configuration. Compare Figures 111-69 and 111-70 (H2) versus Figures HI-74 and 111-75

(CH4). Note that on the methane, Figures 111-74 and IH-75, that the maximum strains peaks out

during the transient while on the hydrogen (Figures 111-69 and 111-70) the highest strains are

very close to the steady-state values. To help understand the difference between the hydrogen

and methane strains aft of the second core cavity, two uncooled analyses were conducted for this

location: one analysis used H2 and 02 combustion products and the other analysis used CH4 and

02 combustion productions. Figures 111-78 and IH-79 show the results of these analyses. Note

that the worst strains occur during the transient with no strain at steady state. Comparing

Figures 111-69, 111-74, 111-78, and 111-79, the cooled methane configuration looks more like the

uncooled curves than the hydrogen cooled configuration for the thick walled section aft the

second core cavity. The benefit of methane cooling was a reduction of strain range from 0.859%

uncooled to 0.82% cooled, while the benefit of hydrogen cooling was to reduce strain range from

0.973% uncooled to 0.684% cooled. Thus, thicker wall sections should be minimized for methane

cooled parts.

The results of this study are based on detailed transient thermal and strain analysis. The

nodal configurations simulating different locations on the foil are provided in Figures 111-80

through 111-83.

The thermal strain ranges of Figures 111-67 through 111-76 were converted to cyclic life

using Figure 111-84. The summary is presented in Figure 111-77. Metal temperatures at full power

are given in Figures 111-85 and 111-86.

2. Conclusions of Preliminary Life Analysis

•' Hydrogen and cooling provides good cyclic life for both thin (0.055 inch) and

thick wall (0.11 inch) sections.

• Methane cooling provides good cyclic life for thin wall sections (0.55 inch).

But for thick wall sections (0.11 inch), methane cooling shows poor cyclic life

improvement.

• Both the methane and hydrogen cooled foils will exceed the 100 cycle life

requirement. (Lives were obtained using nominal material properties.)

111-67
Pratt & Whitney

FR-19009-1

The analysis of the cooled rig blade identified the thick section aft of the rear cavity as

being the life limiting region in the blade. A study was performed to identify the improvement in

life that can be obtained by extending the rear core cavity further back. Figure 111-87 shows the

two cross sections that were analyzed. Section A-A is 0.139 inch thick and is located aft of the

rear core, which has a trailing edge diameter of 0.035 inch. Section B-B is 0.11 inch thick and is

located aft of the extended core, which has a trailing edge diameter of 0.010 inch. A detailed

thermal/strain analysis of the two different configurations showed the following results:

0.139 Inch Section 0.11 Inch Section

CH4 Cooled 0.912%/108 cycles 0.784%/260 cycles

H2 Cooled 0.743%/345 cycles , 0.656%/730 cycles

(Strain Range/Cyclic

Life)

By extending the rear core further aft, the cyclic life of the part was increased by a factor of

more than two.

Two methods of cooling the blade tip cap were considered, as shown in Figure 111-88. A film

cooled design was investigated. Using 0.5% flow* (25% of the entire blade flow) to cool the tip

cap, the film would exit out of 18 0.006-inch diameter holes equally spaced around the tip cap.

This design was deemed unacceptable due to the large quantity of flow and the resultant poor

film coverage.

A convectively cooled design was also investigated. The same type of grooves used to cool

the air flow was used to cool the tip cap, using a cooling flow of only 0.16%. Unlike the film

cooled design, the convective design provides positive cooling capability. The film cooled design

could be a source of problems, if the hot gas secondary flows wash the film off the surface of the

foil. Conversely, this condition does not affect the convectively cooled design.

For this reason, and the reduced cooling flow requirement, the convective cooling tip cap

design was chosen over the film cooled design. The convective cooling tip cap design will have

slightly lower temperatures and strains than the rest of the blade, since the hot gas temperature

should be slightly lower along the rig endwalls than in the middle of the rig.

The nodal breakup for the low cycle fatigue life study of the uncooled slave airfoils is shown

in Figure 111-89. The same hot gas transient used to analyze the cooled foil was used in this

analysis. The maximum strain range was found to be 1.58% on the suction side, Node 192. This

strain range is far in excess of our LCF data for PWA 1447 material. Using a gross extrapolation

of the LCF curve, the cycle life of the slave airfoil was found to be less than 10 cycles.

* Note: Percent flow is defined according to the following equation

Coolant Flow = (Total Rig Inlet Flow) x (% Flow)

O

where flows are in pounds per second and 3 is the number of foils in the rig.

111-68
Core

0.5

0.4

0.3

Thermal

Strain

Range - %

0.2

0.1

Pratt & Whitney

FR-19009-1

Hot Gas Side

a at CD

A = 0.035

B = 0.045

C = 0.023

Hydrogen

^^ ^^ >— ^o' Colad Sid <>iaee

^*^ ^ Hot Side /

Methane

Hot Side

0.06 0.07 0.08 0.09

Overall Wall Thickness - in.

Figure III-63. Thermal Strain Versus Overall Wall Thickness for Rig Conditions

0.10

FDA 301940

111-69
Pratt & Whitney

FR-19009-1

Chordwise Coolant Geometry

Coolant Feed Cavity

Rear Cavity Added to

Minimize Wall Thickness, and

Thus Minimize Thermal

Strains Also Can Be

Used As Intermediate

Leaching Port If Necessary

FDA 301941

Figure 111-64. Rig Blade Coolant Geometry

The information obtained from the previously discussed design studies was used to generate

the geometry of the advanced convectively cooled airfoils and their associated hardware. Even

though the fluidized bed test had not yet been performed, a decision was made to select the

optimum coolant groove-land geometry combination based on the information available. This

decision was made with the understanding that modifications to the drawings could be made

prior to the hardware fabrication commitment if the fluidized bed test results warranted it. The

coolant groove geometry selected for the engineering drawing was 0.035 ± 0.002 inch wide by

0.012 ± 0.002 inch deep. The selected spacing between the slots was 0.035 inch wide nominally.

Table III-9 lists the part numbers and titles of the engineering layout and drawings for the

advanced convectively cooled hardware. This table also lists the figure numbers for these

drawings, which are contained in Appendix B.

F. TASK 4 — AIRFOIL FABRICATION AND TASK 5 — TEST FOLLOW UP

Due to the inability to successfully fabricate the fluidized bed test cylinders, the airfoil

fabrication, test and test follow-up were not performed.

111-70
Pratt & Whitney

FR-19009-1

Rig Gas

Temperature - °F

1800

1600

1400

1200

1000

800

600

400

200

0

Start

Shutdown

Data from Run 117

on 15 December 1981

Run 2 Out of 5

1.0 2.0

Time - Sec

3.0

FDA 301942

Figure 7/7-65. Rig Transient Hot Gas Temperatures

2400

2200

2000

1800

1600

1400

1200

1000

800

600

400

200

0

Rig

Pressure -

psia

Start

Shutdown

'

Data from Run 117

on 15 December 1981

Run 2 Out of 5

1.0

Figure 777-66. Rig Transient Hot Gas Pressure

2.0

Time - Sec

3.0

FDA 301943

111-71
Pratt & Whitney

FR-19009-1

Table 7/7-8. Hydrogen and Methane Configurations At Full Power (Thoi &oa = 1700°F)

Heat Transfer Film Coefficients (h) are in Btu/hr ft? °F

T

'coolant

h

hotgai"coolan t (°F)

Hydrogen Configuration

SS, 30% Axial Chord 15,500 11,000 150

PS, 30% Axial Chord 10,100 11,000 150

SS, 86% Axial Chord 15,300 11,000 266

PS, 75% Axial Chord 11,400 11,000 266

Trailing Edge 13,600 11,000 337

Methane Configuration*

SS, 30% Axial Chord 9,800 2,700 210

PS, 30% Axial Chord 6,400 2,700 210

SS, 86% Axial Chord 9,700 2,700 417

PS, 75% Axial Chord 7,200 2,700 417

Trailing Edge 8,600 2,700 542

•For the methane configuration, we plan on simulating the

CH4+02 hot gas with H2O2 run at lower temperature to give

the same heat flux.

111-72
+0.40

+0.30

+0.20

Thermal

Strain - % +0.10

-0.10

-0.20

Start

Pratt & Whitney

FR-19009-1

Strain Range = 0.427%

Node 192

(Reference Figure III-80)

Shutdown

400 800

Metal Temperature - °F

1200 1600

FOA 301944

Figure 7/7-67. H2 Cooled ATS Blade Thermal Strain Versus Metal Temperature for Worst

Node on Suction Side, 30% Axial Chord

111-73
+0.40 I—

+0.30

+0.20

Thermal

Strain - % +0.10

-0.10

-0.20

Start

Pratt & Whitney

FR-19009-1

Strain Range = 0.380%

Node 192

(Reference Figure III-80)

Shutdown

400 800

Metal Temperature - °F

1200 1600

FDA 301945

Figure 111-68. H2 Cooled ATS Blade Thermal Strain Versus Metal Temperature for Worst

Node on Pressure Side, 30% Axial Chord

111-74
+0.20 I—

+0.10

-0.10 —

Thermal

Strain - % -0.20 —

-0.30 —

-0.40 —

-0.50 —

-0.60

Pratt & Whitney

FR-19009-1

Strain Range = 0.684 %

Node 288

(Reference Figure 111-81)

Steady

State

400 '800

Metal Temperature - °F

1200 1600

FDA 301946

Figure 7/7-69. H2 Cooled ATS Blade Thermal Strain Versus Metal Temperature for Worst

Node on Suction Side, 86% Axial Chord

IH-75
Pratt & Whitney

FR-19009-1

+0.20 I—

+0.10

-0.10

Thermal

Strain - % -0.20 —

-0.30 —

-0.40 —

-0.50

-0.60

Strain Range = 0.652%

Node 338

(Reference Figure 111-81)

400 800

Metal Temperature - °F

1200

Steady

State

1600

FDA 301947

Figure IJI-70. H2 Cooled ATS Blade Thermal Strain Versus Metal Temperature for Worst

Node on Pressure Side, 75% Axial Chord

IH-76
+0.20

+0.10

-0.10

Thermal

Strain - %

-0.20

-0.30

-0.40

-0.50

Pratt & Whitney

FR-19009-1

Strain Range = 0.542%

Node 168

(Reference Figure III-82)

Start

400 800

Metal Temperature - °F

1200 1600

FDA 301948

Figure 111-71. H2 Cooled ATS Blade Thermal Strain Versus Metal Temperature for Worst

Node on Trailing Edge

IH-77
+0.40 I—

+0.30 —

+0.20 —

Strain

Range - % +0.10 —

-0.10 —

-0.20

Pratt & Whitney

FR-19009-1

Strain Range = 0.372%

Node 192

(Reference Figure III-80)

400 800

Metal Temperature - °F

1200 1600

FDA 301949

Figure 7/7-72. CH4 Cooled ATS Blade Thermal Strain Versus Metal Temperature for

Worst Node on Suction Side, 30% Axial Chord

111-78
Strain

Range - %

+ 0.30 I—

0.20

0.10

- 0.10

Start

-400

Pratt & Whitney

FR-19009-1

Strain Range = 0.338%

Node 192

(Reference Figure III-80)

Shutdown

I I

400 800

Metal Temperature - °F

1200 1600

FDA 301950

Figure 7/7-73. CH4 Cooled ATS Blade Thermal Strain Versus Metal Temperature for

Worst Node on Pressure Side, 30% Axial Chord

111-79
+ 0.30 I—

+ 0.20 —

0.10 —

Strain

Range - % - 0.10 —

- 0.20

- 0.30 —

- 0.40

- 0.50

-400

Pratt & Whitney

FR-19009-1

Strain Range = 0.82%

Node 295

(Reference Figure 111-81)

400 800 * 1200

Metal Temperature - °F

1600 2000

FOA 302501

Figure 111-74. CH4 Cooled ATS Blade Thermal Strain Versus Metal Temperature for

Worst Node on Suction Side, 86% Axial Chord

HI-80
Pratt & Whitney

FR-19009-1

+ 0.30 I—

0.20

+ 0.10

Strain

Range - % - 0.10 —

- 0.20 —

- 0.30 —

- 0.40 —

- 0.50

-400

Strain Range = 0.800%

Node 331

(Reference Figure 111-81)

400 800 1200

Metal Temperature - °F

1600 2000

FDA 302502

Figure 111-75. CH4 Cooled ATS Blade Thermal Strain Versus Metal Temperature for

Worst Node on Pressure Side, 75% Axial Chord

111-81
+ 0.30 I—

+ 0.20 —

0.10 —

Strain

Range - %

- 0.10 —

- 0.20 —

- 0.30 —

- 0.40

-400

Pratt & Whitney

FR-19009-1

Strain Range = 0.521%

Node 175

(Reference Figure III-82)

400 800

Metal Temperature - °F

1200 1600

FDA 302503

Figure 7/7-76. CH4 Cooled ATS Blade Thermal Strain Versus Metal Temperature for

Worst Node on Trailing Edge

HI-82
HYDROGEN CONFIGURATION

Pratt & Whitney

FR-19009-1

N = 8000 cycles

CT = 0.427%

T = 220°F

max

Thickness - 0.55 in.

N = > 10,000

tr = 0.380%

T = 210°F

max

N = 750

tr = 0.652%

T - 1380°F

max

N = 560

IT = 0.684%

T = 1450°F max

0.07 in. Thick

N = 2000

fr = 0.542%

T = 1450°F

max

T =1700 to 150°F

70°F

METHANE CONFIGURATION

N = > 10,000

(T = 0.372%

T = 610°F

Thickness = 0.55 in.

N = > 10,000

(T = 0.338%

T = 560°F

T = 1400°F

max

N = 200

(T = 0.82%

T = 1460°F

0.07 in. Thick

N = 2500

tr = 0.521%

T = 1460°F max

FDA 302504

Figure III-77. Cyclic Life Summary

111-83
0.50

+ 0.40

0.30

0.20

+ 0.10

Strain

Range - %

- 0.10

- 0.20

- 0.30

- 0.40

- 0.50

-400

Pratt & Whitney

FR-19009-1

Strain Range = 0.973%

Node 316

(Reference Figure III-83)

Suction Side

86% Axial Chord

I I I I I I

400 800 1200

Metal Temperature - °F

1600 2000

FDA 302505

Figure III-78. Uncookd ATS Blade Thermal Strain Versus Metal Temperature — H2 +

02 Combustion Products

111-84
+ 0.40 I—

+ 0.30 —

-i- 0.20 —

0.10 —

Strain

Range - %

- 0.10 —

- 0.20 —

- 0.30 —

- 0.40

-400

Pratt & Whitney

FR-19009-1

Strain Range = 0.859%

Node 316

(Reference Rgure III-83)

Suction Side

86% Axial Chord

400 800 1200

Metal Temperature - °F

1600 2000

FDA 302506

Figure 7/7-79. Uncooled ATS Blade Thermal Strain Versus Metal Temperature — CHj +

O2 Combustion Products

111-85
Pratt & Whitney

FR-19009-1

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111-89
10

Strain

Range - %

Pratt & Whitney

FR-19009-1

PWA 1447 Thermomechanical Fatigue

Cycle I, 800 to 1700°F

Typical Properties: € _\^ Temperature

TMJ, <; 1900°F

' •*• —. .^ A Extrapolation

T = 1700°F

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Figure 111-84. PWA 1447 Thermomechanical Fatigue

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Rear Cavity

Rear Cavity

Extended

FDA 302514

Figure 7/7-87. Advanced Turbine Study Blade With Extended Rear Cavity

Correctively Cooled

Tip Cap

FDA 302515

Figure IH-88. Advanced Turbine Study Blade Tip Cap Cooling Methods Investigated

111-93
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FR-19009-1

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111-94
Pratt & Whitney

FR-19009-1

Table 7/7-9. Advanced Convectively Cooled Hardware Drawings

Part/No. Title Figure No.

L-238386 L/0 Cooled Turbine Blade Test Rig B-l

T2185434 Cover — Assembly B-2

T2185435 Blade — Assembly B-3

T2185436 Blade B-4

T2185437 Seal B-5

T2185440 Seal B-6

T2184088 Gasket 0.185 X 0.375 X 0.062 B-7

T2184089 Screw-Machine. 0.164-32 X 1.500. Fillister B-8

111-95
Pratt & Whitney

FR-19009-1

SECTION IV

TECHNICAL ACTIVITY — REDIRECTED PROGRAM

The original Advanced Turbine Study (ATS) Program consisted of an advanced convective

cooling configuration as a means of improving the turbine horsepower output at elevated

temperatures for high-pressure reusable rocket applications. This type of cooling configuration

features external cooling grooves on the turbine airfoil surface, which is then coated with a thin

cover. The thin cover conducts heat to the coolant and minimizes the high thermal strains

encountered in the severe heat flux environment of a rocket turbine.

Fabrication difficulties encountered during the sputtering .of the external cover indicated

that an extensive development program was required to execute this cooling design. A program of

this magnitude was beyond the scope of the original Advanced Turbine Study Program. Studies

also showed small or no horsepower payoff, for a hydrogen cycle, at elevated turbine inlet

temperatures and constant airfoil life using this type cooling.

Although this type cooling is still considered the best way to efficiently cool rocket turbine

airfoils without severely sacrificing horsepower output, study results indicated that a program

redirection was required.

Pratt & Whitney recommended that the Advanced Turbine Study Program be redirected to

fabricate a hot core type blade which would provide improved turbine blade life over a

conventional solid blade. This type hollow blade, with hot gas flowing through it, could be tested

in the NASA facility to quantify its life improvement. This improved life could then be traded for

increased turbine inlet temperature (TIT) for higher output horsepower at a given life

requirement, or the improved life could be traded for reduced replacement cost for a given TIT

and mission requirement.

The life improvement and/or the increase in TIT capability for the hot core blade was

derived from a reduction in thermal strain. The thermal strain in the blade was reduced by

thinning the blade wall thickness, thus reducing the thermal gradient across the blade walls. This

was achieved by coring the blades. By flowing mainstream flow through the core cavity a further

reduction in thermal gradient can be achieved, thus creating reduction in thermal strain and an

increase in life and/or TIT capability.

The hot core blade design fabricated from PWA 1480 single-crystal (SC) material

(Reference Figure II-3, Section II) shows a predicted potential of 3.5 times life improvement over

a solid blade design. For the sake of simplicity and cost savings, P&W manufactured these blades

to the same configuration (no twist and no taper) as the previously designed Advanced Turbine

Study blades.

A. TASK 1 — HARDWARE DESIGN

Task 1 consisted of the design of the hot core blade (Reference Figure H-3) and-the

generation of engineering drawings which were used to fabricate the blades and blade holding

fixture.

To relate the life prediction system to actual test results obtained in the Space Shuttle

Main Engine (SSME) blade tester rig, an analysis of the root section of the SSME Ist-stage

turbine blade was performed. In this analysis, the thick cross section near the blade ID was

broken up into nodes as shown in Figure IV-1. A transient analysis was conducted using the

blade tester rig transient recorded on 15 December 1981, run 117, the second cycle out of five.

Since the SSME blade material, Mar-M 246 directionally solidified (DS), is not built into P&W's.

IV-1
Pratt & Whitney

FR-19009-1

material library, a similar nickel-base alloy, PWA 1422, was used for this analysis. As shown in

the table of Figure IV-1, the material properties are quite similar. Results of the analysis yielded

a thermal strain range of 1.59% maximum occurring on the convex surface. Using the low cycle

fatigue (LCF) data for Mar-M 246 DS as shown in Figure IV-2, the cyclic life of the SSME blade

is predicted to be less than 100 cycles. Note that the data in Figure IV-2 is for transversely loaded

(LCF) specimens. This is the weak direction for directionally solidified material. Thus, we would

expect the airfoil cracking to be in the radial direction. The prediction of less than 100 cycles

agrees fairly well with rig test results. It must be remembered that the prediction uses the general

isothermal LCF data and not specific thermal mechanical fatigue data which simulates the exact

strain-temperature history of the rig part. Therefore, a more accurate prediction can not be made

with the available data.

Next, a detailed analysis of the solid slave foil designed for the original programs (but

fabricated out of PWA 1480 material) was made at the same rig conditions. The strain range was

found to be 1.59%. The nodal breakup of this foil is shown in Figure IV-3. The maximum

thickness of this foil closely resembles the shuttle blade root section and consequently has a

similar thermal response. The life of the PWA 1480 solid slave will be similar to the shuttle blade

based on the limited amount of transverse data, as shown in Figure IV-2. However, the data of

Figure IV-4 shows PWA 1480 distinctly better than Mar-M 246 DS and SC with transverse

loading at a lower temperature (1400°F). But the higher temperature'data (Figure IV-2) will more

closely match the rig cycle with maximum temperature of 1700°F.

To improve the LCF life of the solid foil, the part was cored and a study was made where

various levels of mainstream flow were passed through the foil. The following results were

obtained: j

No internal flow — 1.23% strain range

0.52% flow/blade row — 1.18% strain range :

1.17% flow/blade row — 1.17% strain range. !

i

At 0.52% flow, the strain was reduced from 1.23% for the cored, no flow case, to 1.18%.

However, the strain barely changed by doubling the flow. Thus, the 0.52% flow case was selected

as the best choice for reducing the thermal strains on a solid foil. Referring to Figure IV-2, the

cyclic life improvement is shown below. j

Solid Foil 1.59% Strain 64 Cycles

Cored Foil With 1.18% Strain 225 Cycles

0.52% Hot Flow

Life Improvement 225 = 3.5 times

~64

Based on the previously discussed studies, three different blade configurations were

designed: a solid blade, a cored hollow, blade, and a cored hollow blade with an orifice in its

pressure side (hollow blade with a hole). The orifice was sized to allow 0.52% of mainstream flow

(per blade row) to pass through the blade core.

IV-2
Pratt & Whitney

FR-19009-1

IV-3
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FR-19009-1

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THIS IS FROM FINAL REPORT NAS8-33561

GH2 at 34.5 MPa, Freq = 0.067 Hz, R, = -1

Pratt & Whitney

FR-19009-1

Typical Test

Cycle Cyclic

Strain

Inelastic Total

Alloy

Longitudinal

PWA 1480

Transverse

PWA 1480

Transverse DS

MAR-M-246 + Hf

Transverse SC

MAR-M-246 + Hf

Longitudinal SC

MAR-M-246 + Hf

I I II

102

iq3

Cycles to Failure

104 10s

FD 218696

Figure IV-4. Comparison of PWA 1480 (Longitudinal and Transverse) Transverse Single

Crystal and Directionally Solidified, Mar-M 246 + Hf and Longitudinal

Single Crystal, Mar-M 246 + Hf in GH2 at 34.5 MPa, Frequency = 0.067

Hz, R = -1, Temperature = 760°C (1400°F)

The P&W layout drawing for the ATS is shown in Appendix B, Figure B-l (Drawing

No. L-238386). This layout was originally used in the design of the advanced, convectively cooled

turbine blade, and was updated to reflect the hot core blade design. These changes are designated

as change "A" in the Alteration column of the drawing, and mainly consist of the addition of

items 12 through 17, with related views and information. These items reflect the hot core blade

design with and without the pressure side orifice, the solid blade, and the updated blade holding

fixture.

Engineering drawings for the hot core blade hardware were generated and released to

production. Table IV-1 is a list of these engineering drawing numbers (part numbers),

descriptions, and corresponding figure numbers in Appendix B.

IV-6
Pratt & Whitney

FR-19009-1

Table IV-1. Hot Core Blade Hardware Drawings

P&W Part Number

T2185593

T2185592

T2185594

T2185437

T2184088

T2184089

Description

Cover — Assembly of

Blade (Hollow)

Blade (Solid)

Seal

Gasket

Screw — Machine

Figure No.

Figure B-9

Figure B-10

Figure B-ll

Figure B-5

Figure B-7

Figure B-8

Figure B-9 shows the cover assembly (P/N T2185593) of the blade holding fixture. A dash

("—") drawing system is used to identify the different detail parts in each assembly. The parts

list is shown in the lower righthand corner of the drawing, above the title block. The dash one

part number (T2185593-01) identifies the cover assembly, the dash two part number (T2185593-

02) identifies the cover detail, and the dash three part number (T2185593-03) identifies the plug

that is brazed to the cover detail to form the cover assembly. The cover assembly is designed in

two parts to facilitate its fabrication using the wire electrical discharge machining (EDM)

method.

Figure B-10 illustrates the hollow blade assembly (P/N T2185592). The dash one part

number (T2188592-01) designates the hollow blade with an orifice in the pressure side, and the

dash two part number (T2188592-02) shows the hollow blade without the pressure side orifice.

Figure B-ll shows the solid blade assembly.

Figure B-5 depicts the seal (P/N T2185437) used to connect the cavity in the instrumented

blade to the blade holding fixture. The layout drawing for the Advanced Turbine Study

(Figure B-l) shows the assembly of the blades to the cover.

Figure B-8 illustrates the machine screws (P/N 2184089) used to secure the turbine blades

to the cover fixture. Figure B-7 shows the gasket (P/N 2184088) originally used under the head of

the screws (shown in Figure B-8). |

B. TASK 2 — HARDWARE FABRICATION

Task 2 consists of the procurement of raw materials and the fabrication and instrumentation of the hot core blade hardware.

Production hot core airfoils would be fabricated by casting. Because of the costs associated

with the tooling required for airfoil casts, it was decided to fabricate the hot cored blades by

electrical discharge machining (EDM) them from blocks of PWA 1480 single-crystal material.

This was followed by electrical chemical machining (ECM) to remove the EDM recast layer.

Figure IV-5 shows the hot core blade blanks being wire EDM'd out a PWA 1480 single-crystal

material block. ,

The blade holding fixture was fabricated from a cylindrical section of AlSl 347 SST

material. The fixture was turned, milled, and drilled. The blade holding pockets were then wire

EDM'd. Subsequent to the wire EDM of the pockets, previously wired EDM plugs were brazed in

place to form the outer (bottom) face of the blade holding fixture. Figure IV-6 depicts the blade

holding fixture at an intermediate step during the manufacturing.

IV-7

0784M
ORIGINAL PASS IS

POOR QUALITY

Pratt & Whitney

FR-19009-1

FE 357810-4

Figure IV-5. Wire EDM of ATS Hot Core Blade from PWA 1480 Single-Crystal Material

Block

KFAE 347763

Figure IV-6. Cover Assembly of Blade Holding Fixture

To determine the proper operating parameters for the EDM equipment, several solid blades

were machined out of IN100 material. Some of these blades are shown in Figure IV-7. The blade

on the left is a wire EDM blank, machined to the blade root dimensions, i.e., EDM'd out of a

block of raw material. The airfoil shape is then produced by two passes using the plunger EDM

IV-8
Pratt & Whitney

ORIGINAL PASS ^ FR-19009-i

OF POOR QUALITY

method: one pass for a rough cut, with approximately 0.006 inch of overstock (blade shown in

center), and a final finishing pass (blade shown on right). The striations shown on the platform

leading edge of the blade will not exist in a deliverable part. A deliverable part will also have a

more gradual blending of the airfoil portion into the blade root than the blades shown in

Figure IV-7.

FE 357596

Figure IV-7. . Electrical Discharge Machined Blades (IN100 Material)

Figure IV-8 illustrates the three different blade configurations: solid blade (P/N T2185594),

hollow blade without a hole (P/N T2185592-02), and hollow blade with a hole (P/N T2185582-

01). The hollow blade without a hole has 0.055 inch wall thickness. A similar part (P/N

CKD 2078, Figure B-12 in Appendix B) with a 0.035 inch wall thickness was fabricated and

delivered to NASA for testing. The third blade configuration has a 0.050 inch diameter hole in

the blade pressure side near the root, and has the same wall thickness as the blade without the

hole. Removing blade cores reduces the thermal gradient within the blades, therefore reducing

the thermal strain. By locating a hole on the pressure side of the blade, the temperature gradient

is further reduced by allowing hot mainstream gasses to flow through the blade core. The

resulting strain reduction allows the turbine blade life to be extended, or it can be traded for an

increase in TIT. A similar effect is achieved by reducing the blade's wall thickness.

The completed blade holder, shown in Figure IV-9, was designed to fit the existing NASA

blade tester rig and to support the ATS blade without using expensive fir tree or dovetail

attachments. Slots were cut into the holder wall, near the trailing edge that separates the center

pocket from the two outer pockets. These slots, which run the full depth of the pockets, provide

thermal relief of the part during exposure to severe thermal cycles.

IV-9
ORIGINAL P&=2 E3

OF, POOR QUALITY

Pratt & Whitney

FR-19009-1

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IV-10
Pratt & Whitney

FR-19009-1

FE 358184

Figure IV-9. Advanced Turbine Study Blade Holder

The ATS blades installed in the blade holder assembly are shown in Figure IV-10. Two

hollow blades (0.055 inch wall thickness) — one with a hole (center) and one without a hole (left)

— and one solid blade (right) were installed in the holder. The center blade was instrumented

with a thermocouple as evidenced by the thermocouple lead shown in the figure.

Six of the 13 blades fabricated and delivered to NASA were instrumented with

chromel/alumel thermocouples. The thermocouples were placed on the pressure side of the core

cavity wall, approximately Vs of the blade span away from the tip of the blades. Table IV-2

provides the serial numbers of all blades delivered to NASA, the type of blades, and if they were

instrumented.

Prior to delivery to NASA, the blades underwent grain etch and Laiie X-ray to verify grain

quality and grain primary orientation. Appendix C contains the Materials Laboratory report

covering this evaluation.

Upon the completion of the laboratory evaluation, the blades, blade holding fixture, and

associated hardware were delivered to NASA for testing in the SSME blade tester rig. Table IV-3

lists the hardware delivered to NASA.

IV-11
Pratt & Whitney

FR-19009-1

FE 231340

Figure IV-10. Blade Holder With Advanced Turbine Study Blades Installed

Table IV-2. ATS Blades Serial Numbers

Serial No.

1

2

3

4

5

6

7

8

9

10

11

12

13

* Instrumented

Part No.

T2 185592-02

T2 185594

T2 185594

T2185592-02

T2185592-02

T2185592-01

T2185592-01

T2185592-01

T2 185594

CKD 2078

CKD 2078

T2 185594

CKD 2078

Type

Hollow Without Hole*

Solid

Solid

Hollow Without Hole*

Hollow Without Hole

Hollow With Hole*

Hollow With Hole*

Hollow With Hole

Solid

Hollow Without Hole, Thin Wall

Hollow Without Hole, Thin Wall*

Solid

Hollow Without Hole, Thin Wall*

IV-12
Pratt & Whitney

FR-19009-1

Table IV-3. ATS Hardware Delivered to NASA

Part No.

T2 185593

T2185592-01

T2 185592-02

T2 185594

T2 185437

T2184089

T2184088

CKD 2087

CKD 2078

Part Name

Cover, Assy of

Blade (Hollow With Hole)

Blade (Hollow Without Hole)

Blade (Solid)

Seal

Screw Machine

Gasket

Gasket, Washer

Blade (Hollow Without Hole, Thin)

Quantity

1

3

3

4

10

30+6=36'

30

30*

3

* See Task 3, Turbine Blade Testing and Data Analysis,

for section on leak repair of blade holding fixture.

3959C

C. TASK 3 — TURBINE BLADE TESTING AND DATA ANALYSIS

Task 3 consists of conducting the turbine blade tests at NASA and the evaluation of the

test results. The hot core blade concept test program that was submitted to NASA is repeated

below.

Hot Core Blade Test Program

The test rig thermal cycle should provide a nominal chamber temperature of

1700°F and a nominal chamber pressure of 2500 psi at the maximum operating

condition. At the minimum operating condition, the nominal chamber temperature should be —350°F and the nominal chamber pressure should be 160 psi. An

excursion from the minimum operating condition to the maximum operating

condition and back to minimum operation condition constitutes a cycle. The fuel

should be H2, and 02 is used as the oxidizer. Each thermal cycle should take

approximately 16 seconds, or 80 seconds for a five cycle test. The above

conditions were obtained from a memorandum to distribution, dated

15 April 1982, from Mr. Dumbacher (subject: Summary of SSME Turbine Blade

Coating Test Program) and should reflect current rig operating capability and

practice. This memorandum was provided to P&W by Mr. Rex Bailey.

Two hollow blades (one with a pressure side orifice and one without) and a solid

blade will be installed in the holding fixture provided by P&W. The blade in the

center position will be instrumented with a thermocouple, which must be

attached to suitable recording equipment. All standard test rig measurements

(e.g., test rig chamber pressure, chamber temperature, propellant flow rates, etc.)

should also be recorded. The blades should be cycled through the following

sequence:

1. 5 cycles, inspect visually, record any distress

2. 5 cycles, inspect visually, record any distress

3. 5 cycles, inspect visually, record any distress

4. 10 cycles, fluorescent penetrant inspection, record any

distress, photograph

IV-13
Pratt & Whitney

FR-19009-1

5. 25 cycles, inspect visually, record any distress

6. 25 cycles, fluorescent penetrant inspection, record any

distress

7. After 75 cycles, continue testing until all articles have failed

or a maximum of 225 cycles.

The above testing sequence is subject to change, depending upon intermediate

test results and inspection findings.

If during inspection period, a test blade exhibits sufficient distress such that it is

determined that the blade will not survive the test until the next inspection

interval, that blade should be removed and replaced with one of the additional

solid blades provided by P&W.

In order for NASA to provide the necessary recording equipment, a description

of the instrumentation for the hot core blade is provided. The two hollow hot

core blades will each be instrumented with two chromel/alumel thermocouples.

Each of the thermocouple leads will be terminated with standard female Marling

connectors.

During ATS hardware fit and leak checks on NASA's blade tester rig, combustion products

were found leaking through the blade holder assembly. The hardware leaked at the boltheads

that secure the blades to the blade holder, at the orifice where the thermocouple lead exits the

blade holding fixture, and at a braze joint within the body of the blade holder. The braze joint

was required to fill the cavities caused during the manufacturing process when the airfoil pockets

were wire EDM'd into the blade holding fixture. Plugs (the same material as the blade holder)

were wire EDM'd and brazed in place to fill the outer section of the airfoil pockets. The hardware

was returned to P&W to correct the leak condition.

The reoperation consisted of fusion welding the plugs to the body of blade holding fixture

on the outer surface of the blade holder. This surface was built up an additional 0.085 inch min.

Subsequently the six boltholes were redrilled and respot faced, as shown in Appendix B,

Figure B-13. New washers (Appendix B, Figure B-14) were machined out of soft nickel material

to act as gaskets between the boltheads and the recently machined spot faces. The thermocouple

lead leak was eliminated by fabricating an adapter (Appendix B, Figure B-15). The adapter was

aligned with the instrumentation hole and welded onto the outer face of the blade holding fixture.

The thermocouple lead could then be routed through the small diameter hole in the adapter and

brazed to its long, narrow neck with low temperature braze (for easy removal of the

instrumentation lead). Upon completion of the reoperation, the hardware was leak checked and

returned to MSFC for test.

1. Thermal Cycle Test Results

The Marshall Space Flight Center (MSFC) thermal cycle testing under the Advanced

Turbine Study has enabled P&W to validate its turbine airfoil LCF design system and to

demonstrate a SSME airfoil concept that will exceed the 120* cycle LCF life requirement.

Metallurgical superiority of the P&W single-crystal alloy (PWA 1480) is due, primarily, to the

absence of grain boundaries where thermal fatigue cracks initiate. The absence of carbides in the

PWA 1480 matrix also enhances the LCF improvement. Testing has shown, however, that the

most important influence on turbine airfoil LCF life (or strain range) is configuration.

'P&W life goal is 2X NASA's 60 cycle life goal.

IV-14
Pratt & Whitney

FR-19009-1

On 11 September 1984, testing began at MSFC using P&W airfoil specimens in the NASA

thermal cycle test rig (Figure IV-11) mounted on test stand No. 116. At the time of this writing,

two tests evaluating six P&W airfoil specimens have been conducted. One specimen simulating

the P&W improved SSME high-pressure fuel turbopump (HPFTP) blade design demonstrated a

LCF life in excess of the 60 cycle goal, in the simulated SSME environment, as shown in

Figure IV-12.

Exhaust

P - 160 to 2500 psi

T - -350°F to 1700°F

A - Approximately 6500°/sec

FD 283160

400

0

-400

Figure IV-11. Marshall Space Flight Center Thermal Cycling Rig

20001—

1600

1200

Temperature 80Q

10 20 30 40

Time - sec

50 60 70

FD 283161

Figure IV-12. Thermal Cycling Test Environment

The initial test objective was to establish a baseline using three specimens described in

Table IV-4. All three test specimens, as shown in Figure IV-13, were machined using electrical

discharge machining (EDM) followed by electrochemical machining (ECM) to remove the recast

layer. The calculated strain range of the ECM finish condition specimens from the initial test are

shown in Table IV-4.

IV-15
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Table IV-4. Test No. 1 Specimen Configuration and Strain Range

Specimen No. Wall Thickness (in.) Strain Range (%)

0.190 (solid)

0.055 (hollow)

0.055 (hollow vented)

1.59

1.23

1.18

Solid

ECM Finish

IA1-3*

0.055 in. Wall Thickness

ECM Finish

IA4-5*

0.055 in. Wall

Thickness ECM Finish

0.050 Vent

Hole

IA5-7 *

*Dash Number Corresponds to Blade Serial Numbers in Table IV-2

FD 290874

Figure IV-13. Test 1 PWA 1480 Single-Crystal Material

Specimen No. 1 simulated the current SSME HPFTP Ist-stage turbine blade airfoil shape

and strain range, but substituted PWA 1480 single-crystal for Mar-M 246 DS alloy. Specimens

No. 2 and 3 are cored configurations with reduced wall thickness to 0.055 inch, with the No. 3

blade also demonstrating the advantage of hot gas circulating into the core through a 0.050-inch

diameter hole on the pressure wall near the platform fillet radius.

Test No. 1 progressed with inspections at cycle intervals of 5, 10, 23, 33, 43, 53, 63, and

71 cycles. Microscopic examination (25X), supplemented by black-light photography of the

fluorescent penetrant inspection (FPI) results, was used to evaluate the blades condition. Initial

cracks and crack propagation were carefully mapped.

The second test was conducted on three different specimens, as shown in Figure IV-14.

Airfoil specimens 4 and 5 were similar in configuration to the No. 1 and 2 from the initial test,

but with improved external finishes. Specimen No. 6 contained an enlarged core which thinned

the wall thickness to 0.035 inch, closely simulating the P&W HPFTP blade designs. Calculated

strain ranges are presented in Table IV-5.

IV-16
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Solid

Buff Polish

0.055 in. Wall Thickness

Cast Finish

0.035 in. Wall Thickness

Cast Finish

IB6-2 * IB9-1 * IC2-13 *

*Dash Number Corresponds to Blade Serial Numbers in Table IV-2

Figure IV-14. Test 2 PWA 1480 Single-Crystal Material

Table IV-5. Test No. 2

FD 290875

Specimen No.

4

5

6

Wall Thickness (in.)

0.190 (solid improved finish)

0.055 (hollow improved finish)

0.035 (hollow improved finish)

Strain Range (%)

1.59

1.23

0.94

Specimen No. 4 was identical to No. 1 except the airfoil surface was polished to eliminate

the surface pitting associated with the ECM process. Specimen No. 5 was similar to No. 2 except

its airfoil surfaces were polished and then vapor blasted to remove polishing lines and to simulate

a cast airfoil finish. Specimen No. 6 is the first test airfoil simulating the lower strain range of the

P&W SSME HPFTP designs. Similar to specimen No. 5, it too has a surface finish simulating a

cast airfoil.

Test No. 2 progressed through 73 cycles with inspection intervals at 3,10,11,15, 23, 26, 30,

32, 37, 47, 54, 62, and 73 cycles. Tests 1 and 2 inspection intervals and initial crack determination

are depicted in Figure IV-15. Photographic documentation of the several test configurations is

shown in Figure IV-16. Detailed inspection of the parts was aimed toward initial crack detection

and subsequent propagation. Figures IV-17 and IV-18 depict crack initiation and progression

throughout the two tests. General conclusions from this testing are:

• The preferred P&W thin-wall airfoil design with P&W 1480 single-crystal

material has the potential to meet LCF life goals and double the current

SSME HPFTP blade life

• Initial cracking generally occurred on the pressure side, in the platform fillet

radius

• Cored parts are significantly more crack resistant than solid parts

• Thin wall and ventilated parts are the superior configurations for this rocket

engine environment.

IV-17
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2. Metallurgical Analysis

Completion of two series of tests on specimens made from PWA 1480 (nickel-base singlecrystal alloy) has supported the potential selection of this alloy for use in SSME turbine airfoil

applications. Distress evident on these specimens corresponds to locations of calculated

maximum strain due to thermal stresses. Alloy PWA 1480 has been extensively characterized in

thermal mechanical fatigue (TMF) under simulated turbojet engine conditions. The life of the

engine parts closely matched the life estimated from extrapolation of this data.

The test airfoils had surface connected cracks which have features similar to laboratory

TMF samples. Maximum depth of cracking was seen at the fillet radius, of approximately

0.030 inch, becoming more shallow when approaching the blade tip. Specimen 1A1-3, solid airfoil

specimen simulating the current SSME design, was cut up and extensively evaluated in the P&W

Metallurgical Laboratory. Figure IV-19 shows the condition of the part after 71 cycles of testing,

the location of the blade's section, and defines the LCF crack patterns. Photomicrographs shown

in Figures IV-20 through IV-25 document the extent of crack propagation into the structure of

the airfoil.

Surface finish affects crack initiation. The three test airfoils from Test No. 1 exhibited

surface pitting from the machining and grain etching operations. This surface is shown in

Figure IV-26. The location of the pitting corresponds to the dendrite cores of the single-crystal.

Post-test evaluation showed cracks connecting these pits.

In the second test sequence an attempt was made to improve the surface finish of the parts

by hand polishing. The solid part (No. 4) was polished to a brushed appearance. This finish did

not significantly improve the crack initiation life of the part, but did change the direction of

cracking. The cracks followed the polishing grooves across the span of the airfoil, indicating the

sensitivity of crack alignment to surface finish. Specimens No. 7 and 8 were polished and vapor

blasted to a matte casting surface finish. Tests with these specimens simulating a cast finish

showed that crack initiation was not associated with surface anomalies.

The correlation of hydrogen embrittlement on the life of these parts is incomplete. Analysis

methods used to date have not given any positive indications of embrittlement. Material testing

has indicated that most materials, PWA 1480 included, are least affected by hydrogen

embrittlement at either cryogenic or high temperatures, and that maximum embrittlement

occurs at room temperature (RT). By extrapolating this characteristic to the test condition a

minimal effect of hydrogen will be expected since rapid heating and quench rates make the

hydrogen exposure at RT a small portion of the total cycle duration.

3. Resumption of Testing

When testing resumes, specimens listed in Table IV-6 and identified as Serial Nos. 4, 6, 9,

11, 12, Rl, R2, R3, and R4, will be available for evaluation. Serial Nos. 4 and 9 will define the

degree of LCF improvement with increased substrate densification via HIP processing. Serial

No. 6 will assess the benefit of increased hot gas circulation in the core cavity. The 0.100-inch

diameter hot gas entry hole is aligned with the direction of hot gas flow to increase internal

circulation in the platform region where initial cracking has been observed due to the high

thermal strain associated with the heavier thickness in this area. Serial No. 11 is a duplicate of

the successful blade specimen (Serial No. 13). This particular specimen (Serial No. 11), while

manufactured identical to Serial No. 13, has a grain anomaly that could have a degrading effect

on LCF life. The defect is a "high angle grain boundary" that is documented in the photos shown

in Figure C-8, Appendix C. This part warrants test to evaluate the influence of grain anomalies

on LCF life. Serial Nos. Rl through R4 are duplicates of the different blade designs, but

fabricated out of NASA-supplied Mar-M 246 alloy for material comparison with the PWA 1480

SC material.

IV-22
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FD 290879

Figure IV-26. Surface Pitting on Initial Low Cycle Fatigue Rig Test Specimens

At NASA's request four additional blades were delivered to MSFC under Contract

Modification 12. The material used was Mar-M 246, which was supplied by NASA. The hollow

blade configuration had no vent holes and is described below.

S/N Description

Rl Hollow (0.035 in. wall)

R2 Hollow (0.055 in. wall)

R3 Solid

R4 Solid

Reference P/N

CKD 2078

T2185592-2

T2185594

T2185594

These specimens are also identified in Table IV-6 as Serial Nos. Rl through R4. These test

specimens will provide an excellent correlation between the Rocketdyne SSME alloy and P&W's

single-crystal alloy (PWA 1480).

IV-30
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SECTION V

CONCLUSIONS AND COMMENTS

The difficulties encountered during the fabrication of the closeout layer for the advanced

convectively cooled scheme precluded the evaluation of the original approach for cooling turbine

blades in the high-temperature environment encountered in today's high-pressure reusable liquid

rocket engines. These difficulties consisted of the absence of adhesion between the sputtered

closeout layer and the parent material of the substrate blade. The lack of adhesion was attributed

to the formation of surface oxides or impurities on the surface of the blade.

If these problems could be overcome and an alternative method for applying the closeout

layer (the key to the advanced convectively cooled design) could be developed, the advanced

convectively cooled blade concept would then become a strong potential candidate for increasing

the life and performance of turbine blades in high-pressure liquid rocket applications, especially

for hydrocarbon fueled rocket engines. Due to complexities encountered during the fabrication

process, the advanced convectively cooled blade must be regarded as an area of technology in

need of further basic research.

On the other hand, the use of PWA 1480 advanced single-crystal materials and the hot core

blade design can bring near-term payoffs in liquid rocket turbine blade life and performance. The

benefits of the implementation of these technologies have been clearly demonstrated by the

results of the testing performed with the hardware fabricated in the Advanced Turbine Study

Program.

Optimization of the hot core blade design should be pursued. Hot core blade implementation into an aerodynamically optimized design (twisted and tapered blade) and the evaluation of

its benefits in a real application appear to be the logical steps to follow.

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APPENDIX A

SUMMARY OF SPUTTER DEPOSITION WORK

A. INTRODUCTION

Sputter deposition is one of the physical vapor deposition techniques in which a coating is

built up on a substrate - essentially atom by atom. In a physical sputter deposition system, ions

are accelerated to a target with sufficient energy to knockout or "sputter" atoms from the target

surface. These sputtered atoms travel with considerable velocity until intercepted by the

substrate to build up a coating. The process can be accomplished by a beam of ions (or neutral

atoms) or by accelerating ions from the plasma of a gas discharge which is established near the

target.

Sputter deposition technology has reached a level of maturity that is beyond the laboratory

curiosity stage. Numerous substrate-coating combinations have been demonstrated — particularly elemental or simple compound thin films for electronic and optical applications. Fewer

combinations of thick alloy deposits on engineering material substrates have been investigated.

In either case, work on new combinations proceeds from experience and not by precise modeling

or knowledge of the deposition process. Definitive experiments would require a detailed

characterization of substrate, coating, and electrical discharge which includes:

• Adhesion

• Coating structure, defects, and crystallography (including orientation)

• Composition of substrate surface and coating including trace elements

• Coating stress, thickness, and hardness

• Substrate roughness

• Substrate temperature

• Mechanically worked subsurface layer (Beilby or selvedge layers)

• Mechanical properties of the coating

• Coefficient of thermal expansion for the coating

*

• Chemical solubility of coating and substrate

• Process parameters — plasma characteristics and particle species, energy

and density impacting the substrate.

Obtaining these details involves designing intricate experiments and characterizing the

results with sophisticated techniques. However, this is seldom done due to the considerable

expense involved.

Experience which served as the background for this deposition program included:

• NASA LeRC contracts NASA CR 134824, NASA CR 135153, and NASA

CR 159637, "Development of Sputtered Techniques for Thrust Chambers."

In this program, thick copper, titanium, and nickel-base alloys were

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deposited and characterized. Bulk coating properties were obtained. Copper

alloys were welded. Two thrust chamber throats were fabricated as

illustrated in Figure A-l by sputter deposition of thick Cu-0.15 Zr over

sputter deposited aluminum filler.

• NiCoCrAlY coatings are routinely sputter deposited on turbine airfoils to a

0.005 inch thickness in a system designed specifically for this purpose.

Deposits up to 0.100 inch thick have been deposited in the same equipment.

Forty-eight F100 Ist-stage turbine blades can be coated in a six hour cycle in

this system.

• NiCoCrAlY alloys have been deposited at rates in excess of 0.007 inch per

hour in the hollow target system used in this program.

The next three sections in this summary of sputter deposition will cover the following:

• Selected details on the sputter deposition process

• Deposition work done to date and conclusions reached

• Additional work required to complete the program.

It will be found that many of the technical elements required for this program have been

worked out and that the sputter deposition process is sufficiently versatile to accomplish the

task. No technical impossibilities are known at this time.

B. SPUTTER DEPOSITION— ELEMENTS OF THE PROCESS

Earlier in this report, a sputtering system equipment and some of the deposition process

were described briefly. In this section, the sputter deposition process will be described in more

detail to provide the background for understanding program deposition work to date. The

discussion will be directed particularly to the substrate and those factors that affect adhesion of a

coating to it.

The starting point for substrate phenomena will be the coating growth process. Electron

microscope work on defined single-crystal surfaces has shown that the process proceeds by:

• Atom impact

• Atom lateral movement

• Nucleation . -

• Island growth

• Coalescence of islands

• Thickness buildup.

While these steps cannot be demonstrated precisely on the surface of an engineering

material, they should apply to a sufficient extent to guide modeling of important process

parameters.

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Work with thick deposits has demonstrated that coating structure and growth habit,

indicative of modification of the basic growth process, are influenced by substrate surface

features. These are:

• Surface macro defects such as scratches, pits and protrusions (cleaning

particle grit, diffusional artifacts, dust, etc.)

• Substrate temperature

• Angle between the surface plane and the adatom flux

• Surface composition and crystal structure.

Adverse effects on coatings, such as leaders (through defects) and poor adhesion, are

minimized or eliminated by careful substrate preparation and manipulation of process

parameters, which was discussed in this report.

The basic growth process can also be modified by interaction of energetic particles with the

substrate and/or the growing deposit (a continuously changing substrate). Energetic particles

originating at the sputtering system target include:

• Electrons

• Neutral atoms of target material (the adatoms)

• Reflected neutral sputtering gas atoms

• Photons '

• Negative ions.

Particles originating from the gas discharge (plasma) include:

• Ions of sputtering gas

• Electrons

• Neutral atoms

•• Photons

• Chemical complexes (molecules).

In a gas discharge sputter deposition system, these particles cannot be controlled

independently of each other.

However, these particles can be controlled to a sufficient extent to affect modification of

coating adherence and growth habit. For example:

Increasing Target Voltage

• Increases the average adatom energy

— Increases adhesion (ion mixing, implantation, etc.)

— Heats target which may cause elemental diffusion, phase

changes, and outgassing of the target and radiation heating

of the substrate.

— Sputtering gas burial (pumping)

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• Increases secondary electron energy

— Heats the substrate

— Probably affects nucleation rate

• May increase deposition rate if in range where yield is still increasing

Increasing Negative Substrate Bias Voltage

• Reduces coating thickness buildup rate by re-sputtering

• Heats the substrate

• Modifies coating structure by forward sputtering to fill grooves and

eliminate shadowing

• Cleans or contaminates, depletes alloys of select elements, forms artifacts,

and enhances diffusion depending on substrate temperature, species present,

and duration of bombardment

Increasing the Discharge Pressure*

• Decreases average adatom energy by increasing collisions with neutral gas

atoms

• Modifies electric field details at substrate and target through sheath spacing

changes

Probably affects the nucleation process.

Gas-Induced

Zone 1 Structure

Shadowing

Processes

Surface

Diffusion

Processes

Bulk Diffusion

Processes

Composite

Structure

Processes

Inert Gas

Pressure - Pa

0.6

0.15^,0.2-

Substrate-Induced

Zone 1 Structure

Isolated Zone 1

Growth Defects

Zone 2

Zone T

Zone 3

Recrystallization and

Grain Growth

0.8 Substrate

Temperature - T/Tm

FD 124536

Figure A-2. Coating Structure as a Function

* A phenological description of growth habit as a function of discharge pressure and substrate temperature has been given

by Thornton as illustrated in Figure A-2.

A-5
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One of the prime requisites in any coating work is the adhesion of the coating to the

substrate. Adhesion can be classified as two types — mechanical and chemical. Mechanical

adhesions is an interlocking of the coating with substrate hills and valleys. Chemical adhesion

takes numerous forms with actual interdiffusion of the substrate and coating being an important

mechanism in metallic systems. If mechanical adhesion is disregarded, adhesion will depend on

the chemical composition of the substrate, and in particular, on "contaminants" on the surface.

Oily layers of any thickness preclude adherence. Extremely thin layers of oxides and other stable

compounds and elemental material such as carbon and boron can act as barriers to

interdiffusion. (Stable compounds are not always detrimental and are sometimes deposited

deliberately to aid adherence in many hard coating systems).

Oily films are usually completely eliminated by cleaning the substrate before installation in

the sputtering system. A typical process is scrubbing with deionized water and alcohol rinses.

Improperly designed vacuum pumping systems may replace the layer in the system. Elimination

of diffusion barriers is more difficult since the barrier can be formed during the early stages of the

deposition cycle including the so called "sputter cleaning" cycle. Initial substrate heatup can

cause diffusion of elements from the substrate to or over the surface. Establishing the plasma in

an unbaked deposition system can "scrub" water vapor and other contaminants from all

surfaces. Water vapor can be disassociated into oxygen and hydrogen which are both persistent

in the system. Application of low negative voltages to the substrate can actually form oxide layers

instead of cleaning it. These same layers may be formed on the target and then sputtered off.

Other sources of contamination are virtual leaks (from unvented blind-tapped holes), the

sputtering gas and the target material.

Substrate cleaning in the deposition system can be accomplished by one or all of the

following techniques:

• Heating the substrate

• Ion bombardment with sputtering gas (usually argon or krypton)

• Hydrogen gas mixture to reduce oxides (forming gas)

• Oxygen mixture to gasify carbon to CO and CO2 which can be pumped out of

the system

• Intermittent pump down to ultra high vacuum during one of the above

cleaning processes.

The cleaned surface will not be atomically clean. However, it will be sufficiently clean for an

adherent coating to form.

The proof of surface cleaning can be obtained from testing for adhesion. Typical adhesion

tests include:

• Epoxy stud pull test (good to 10,000 psi on thin films)

• Scratching

• Diamond indent

• Energy pulse (peening, ultrasonic, laser, etc.)

• Shearing

• Bending.

Of course not all of these tests can be applied to every material and quantifying the

adhesion can be difficult. In this program, adhesion must be sufficient to withstand any stress

A-6
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developed during handling, leaching, and if required, peening of the 0.010 inch thick closeout

layer.

If adhesion is not obtained, substrate surfaces must be looked at in detail with surface

sensitive instruments (auger electron spectroscopy, etc.). Correlations are then made with

process parameters.

C. SUMMARY OF DEPOSITION WORK TO DATE

The program was started with the straightforward approach shown schematically in

Figure A-3. Several differently sized grooves are machined in cylindrical specimens. The grooves

are filled and excess filler removed to present a smooth substrate surface to the adatom flux.

After cleaning, the cylinders are installed in the coater, and the PWA 1447 closeout layer is

deposited. The filler material is removed to leave passages in the structure. Optional steps such

as peening of the closeout layer and heat-treating may be required.

Fill Grooves With Removable Filler

—» Sputtered Aluminum

—> SiO2 Slurry

—. Aluminum Wire

• Sputter PWA 1447 Cover

Remove Filler Material

• Heat-Treat To Homogenize If Necessary

FD 267789

Figure A-3. Direct Approach to Fabricating Convectiuely Cooled Structures

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Each step of the general approach was subject to side experiments before a cylindrical

specimen was attempted. The following results were obtained from these experiments:

1. Filler Materials

• Interdiffusion of sputtered aluminum and the PWA 1447

substrate was excessive at substrate temperatures greater

than 900°F. At substrate temperatures below 700°F the

aluminum appeared to be highly stressed and easily debonded from the substrate. A 6061 aluminum target was used and

the substrates were cooled to various degrees by their

proximity to a water-cooled mandrel.

« Although filling of larger grooves with aluminum wire was

satisfactory the wire could not be contained in the shallowest (0.012 inch deep) grooves due to a lack of a vertical side

wall. The radius from the groove comers and the rounding of

the land edges effectively eliminated a vertical wall on these

grooves.

• The colloidal silica lacked both adhesion to the groove walls

and cohesion to itself. This sand-like material was pulled out

by the closeout layer. This type of filler may still have merit

with a different formulation.

2. Sputtered PWA 1447 Closeout Layer

• Deposition under normal conditions in the hollow target

systems caused excessive interdiffusion of the aluminum

filler and closeout layer. Temperature was approximately

1200°F at the substrate.

3. Aluminum Leaching

• Aluminum was leached satisfactorily from a simulated

filled groove. This groove was formed by electroplating

nickel over a 4043 wire which was positioned on a Ni-200

• base material. A sodium hydroxide solution at 150°F with

periodic ultrasonic agitation removed all of the aluminum

from a 1.0 inch length of a 0.035 by 0.012 inch (cross

section) groove in 120 minutes.

4. Closeout Layer Characteristics

• Electron microprobe testing revealed that the closeout layer

material was the same composition (within experimental

error) as the target material.

• Hardness of the as-deposited material was 429 HV which

compared nicely with the 395 HV for the cast PWA 1447

substrate bar.

One cylindrical specimen was filled by a combination of aluminum wire inlay and sputter

deposition. This cylinder was placed over the water-cooled mandrel and rotated during the PWA

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1447 closeout layer deposition. The cylinder had an excellent appearance after deposition but

suffered extensive debonding of the closeout layer during the leaching cycle. Delamination

occurred even at the end portion of the cylinder which was 0.5 inch away from the nearest

grooves. Although the delamination left several large pieces of the debonded closeout layer

material intact, the zone one type structure (essentially through defect network) was evident

(Figure A-3). Even though this type of structure might possibly have been peened (with filler still

in the grooves) to "cure" the defects, cooling actual turbine blades with this type of coating

structure would not be practical.

At this point, another approach was planned as shown in Figure A-4. The heart of the

alternative approach was to have a cylindrical substrate that could be closed out without

temperature limitations imposed by the aluminum filler. The higher temperature during

deposition would benefit both adhesion and structure of the closeout layer. Nickel was chosen for

the intermediate 0.001 inch thick (nominal) layer for two reasons. First, this 0.001 inch layer

must be adherent and defect free. With the end product in mind, a "cold" sputtering process

(Figure A-5) was chosen to deposit this layer. Use of an elemental material would be more

amendable than an alloy to the deposition techniques used to obtain the required properties.

Secondly, a nickel target for this system was available. It was recognized that this nickel layer

would require a structure homogenizing heat treatment schedule.

The results to date of the alternative approach are summarized below:

• Thin nickel films deposited in the RF diode system on Ni-200 and PWA

1447 substrates are adherent to the 10,000 psi limit of the epoxy-bonded

tensile stud test. Thin films are required for this test which precluded testing

the 0.001 inch thick closeout layers by this technique.

• Full thickness Ni closeout layers over aluminum wire-filled PWA 1447

substrates were not adherent. Like the cylindrical specimen described in

Section B, appearance after deposition was excellent, but debonding

occurred during leaching.

• For a side experiment in the hollow target coater, PWA 1447 was deposited

(0.001 inch) onto PWA 1447 and Ni-200 substrates. The substrates were

attached to the water-cooled mandrel to simulate conditions used for the

filled cylinder substrates. Bend tests showed that the deposit was adherent

to the Ni-200, but not to the PWA 1447.

These adherence results point to surface contamination as being the key problem rather

than a difference in coefficient of thermal expansion between the aluminum filler and the PWA

1447 substrate. Calculations have shown that the difference is not sufficient to strain the

closeout layers under the deposition and leaching conditions used.

D. PROGRAM REQUIRED TO COMPLETE SPUTTERING WORK

Evidence to date requires that details of the PWA. 1447 surface be examined to determine

the cause of the lack of adhesion. A program to achieve this is outlined below where the species

remaining on the surface versus the processing parameters will be examined. Also their effect on

adhesion (and possibly on structure) will be determined. A parallel program should be run to

determine the heat treatment schedule required to homogenize the closeout layer(s) and to

obtain bulk (or better than) PWA 1447 mechanical properties. Additional materials would be

required. The program would be expensive and lengthy due to nonavailability of surface

characterizing equipment at P&W.

A-9
Prat

t & Whitney

FR-19009-1

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A-10
Pratt & Whitney

FR-19009-1

a. System Equipment — Planar Target RF Diode Sputtering System

13.56 MHz

Rf potential

•Water-cooled Target Holder

sjs^s?$^^

6 inch Diameter Sputtering Target

1.5

rpm

Substrate

RF Bias

(% of Target Potential)

b. Schematic of Deposition Chamber (Not to Scale)

Figure A-5. Radio Frequency (RF) Diode Sputtering System

FD 267791

A-ll
Pratt & Whitney

FR-19009-1

1. The experimental program for depositing the 0.001 inch layer is as follows:

a. Prepare flat and cylindrical substrates, with and without

grooves (cast, machine, and degrease).

b. Characterize substrate surface finish and composition.

c. Fill grooved substrates with pure aluminum and Al alloy

fillers. Refinish and degrease.

d. Characterize filled substrate surface finish and composition.

e. Clean substrates in the vacuum system — a parameter

study — using both filled and unfilled substrates (not

simultaneously).

f. Characterize clean substrate surface compositions and

structures.

g. Deposit closeout layer over substrate cleaned with selected

parameters — use filled and unfilled substrates.

h. Characterize the deposits for adhesion, structure, topography, and stress.

i. Characterize any debonded material to establish cause of

adhesive failure.

j. Determine effects (if any), of the gas discharge plasma and

deposition parameters and adjust for adhesion as required.

k. Process filled and unfilled cylindrical specimens with best

parameters (clean, closeout, and leach).

1. Overcoat 0.001 inch thick closeout layers with a nominal

0.015 inch of PWA 1447 and machine to 0.010 inch

thickness (peen if necessary). Leach the grooved cylinders,

heat-treat (from parallel program), and fluidize bed test.

2. The parallel program to establish homogenizing heat treatment schedule is

as follows:

a. Deposit 0.001 inch of Ni on PWA 1447 specimens and

overlay with PWA 1447. Machine specimen to have an

even 0.010 inch thick (total) closeout layer. Characterize

as-deposited structures. Peen if necessary.

b. Diffuse with selected heat schedules. Characterize the

heat-treated structures for structure, composition, and

diffusional type defects. Note: If excessive void formation

at the interfaces persists, iterate this program using

specimens without the 0.001 inch Ni layer. Assuming Ni

caused the diffusional defects, change the program at step

l.g to include a 0.001 inch PWA 1447 layer.

A-12
Pratt & Whitney

FR-19009-1

c. With best parameters, sputter deposit, heat-treat and

fabricate mechanical test specimens to compare best

sputtered material mechanical properties with case

material properties.

d. Iterate as necessary (change hollow target composition,

etc.) to optimize properties of the sputtered material.

3. The alternate approaches to this program are as follows:

a. With the closing out of airfoils in mind, investigate

electroformed Ni for the 0.001 inch thick closeout layer. If

this were successful, the filler machining step would be

eliminated in the process. The Camin Lab process may

lend itself to fabricating this layer.

b. Alternate sputtering processes, such as planar magnetron

and post-magnetron sputtering, could be investigated.

c. These processes are intermediate in temperature and may

be used to deposit thick PWA 1447 layers over the

aluminum filler without excessive substrate heating and

interdiffusion of filler with the substrate and closeout

layer. By this same reasoning, sputter deposition, if

required, of the aluminum filler could be accomplished on

an uncooled substrate.

4. Nondestructive quality control techniques for a manufacturing process

should be investigated early in any program.

5. The materials required are as follows:

a. PWA 1447 flat casting for cut up into test specimens

substrates including mechanical test specimen substrates

b. PWA 1447 cast cylinders — plain and for grooving

c. PWA 1447 cast disk for 6 inch diameter target fabrication

d. Modified PWA 1447 cast hollow target.

A-13
Pratt & Whitney

FR-19009-1

APPENDIX B

ADVANCED TURBINE STUDY

ENGINEERING DRAWINGS

B-l
Pratt & Whitney

FR-19009-1

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B-2
Pratt JiWhjtney

FR-19009-1

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B-3
Pratt & Whitney

FR-19009-1

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FR-19009-1

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B-5
Pratt & Whitney

FR-19009-1

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FR-19009-1

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FR-19009-1

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Pratt & Whitney

FR-19009-1

APPENDIX C

VERIFICATION OF ATS BLADE GRAIN QUALITY AND ORIENTATION BY GRAIN ETCH

AND LAUE X-RAY METHODS

C-l
Pratt &..Whitney.

FR-19009-1

PftATTAWHITMEY AIRCRAFT GROUP

GOVERNMENT PRODUCTS DIVISION

& IBCBWMCT

MEMORANDU M REPOR T

ltua w 26133

IR&D

(3126) Department

CHARGE HO. E271-30-200-72

DATE 8-27-84 _

CASTING JOB A19Q

REQUESTED BY J. H. Castro RL10 Engineering

Send To: H. M. Gibson. D. E. Paulus. J. L. Price. T. A. Rackley III, C. C. Rhemer.

U C. Rinfl. r. B. Stevens. D. C. Stewart. D. B. Tipton

Work Requested Perform orain etch and Laiie X-rav to verify grain quality and

PERTINENT DATA

Part Name:

Serial Number:

Material Spec.:

Mfg. Source:

Machining:

RESULTS

SSME Turbine Blade

1-1 3

PWA 1480

PWA Manchester Foundry

GPD

Primary orientation recorded on all thirteen blades was within the commonly

accepted limit of 15°. Three blades (S/N 8, 9, 11) contained low angle

boundaries which would be acceptable while blade S/N 10 contained several

high angle boundaries and was considered unacceptable.

Reported by C. M. Biondo _ _ Approvals

E«. 5273

. Tables: 2 Figures:

. 9

C-2
Pratt & Whitney

FR-19009-1

L/N 26133

REPORT

As part of an advanced turbine study, NASA Contract NAS 833821, PWA 1480

blades machined to the Rocketdyne configuration will be tested in a rig

at Huntsville, AL along with Rocketdyne blades cast in MAR-M-246. The

initial effort consisted of evaluating eight cast PWA 1480 blocks supplied by Manchester Foundry (see Report, ME&T 25608). From the investigation two blocks were selected and thirteen blades (of varying •

designs) were subsequently EDM'd from the blocks at Government Products

Division as shown in Figure 1.

All blades were grain etched and Laiie X-ray was used to verify primary

orientation and measure mismatch on any grain boundaries. Primary

orientation was determined by taking Lau'e X-ray shots on the tip of the

blade as outlined in Figure 2. There is no current standard which specifies quality requirements for this blade configuration, but in all cases

the primary orientation recorded (Table I) was within the commonly accepted

limit of 15°.

Typical grain etched condition of all thirteen blades is shown in Figures

3-9 . Four blades (S/N 8, 9, 10, 11) were observed to contain grain

boundaries which were then Laue X-rayed to determine mismatch between the

specific grain and the parent crystal. Three blades (S/N 8, 9, 11) contained low angle boundaries which would be acceptable to existing

specification limits (CVS-6) while blade S/N 10 contained several high

angle boundaries and was considered unacceptable.

C-3
X-RAY DIFFRACTION RESULTS:

BLADE ORIENTATION

Pratt &..Whitney..

FR-19009-1

((X)

DEGREES OFF AXIS

1

2

3

4

5

6

7

8

g

10

11

12

13

[100]

[100]

[100]

[100]

[100]

[100]

[100]

[100]

[100]

[100]

[100]

[1003

[100]

10

10

3

a

10

g

g

g

11

7

7

B

g

TABLE I

X-ray diffraction data showing measurement

of primary (100) orientation for each test

blade.

L/N 26133

C-4
Pratt & Whitney

FR-19009-1

X-RAY DIFFRACTION RESULTS:

BLADES

B-A. B-B

9- A. 9-B

10-A. 10-C

10-B. 10-C

10-D. 10-E

10-F. 10-6

10-S. 10-H

11-A, 11-B

TILT

MISMATCH

m

1

1

12

2

9

5

2

1

ROTATION

MISMATCH

(R)

0

1

8

3

12

5

1

1

'TOTAL

MISMATCH

1

1

14

A

15

7

2

1

TABLE II

X-ray diffraction data giving measurement of

grain mismatch on blades which contained defects.

Figures 5,6,7 outline the location of Lade

shots and labeling of specific grains to determine

mismatch.

L/N 26133

C-5
Pratt & Whitney

FR-19009-1

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Pratt & Whitney

FR-19009-1

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ORIGINAL PAGE IS

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Pratt & Whitney

FR-19009-1

FAL 79653 S/N 1 Mag 2X

FAL 79655 S/N 2 FAL 79656 S/N 2 Mag 2X

FD 301902

Figure C-3. Grain Etched Condition of Experimental Blades

C-8
Pratt & Whitney

FR-19009-1

FAL 79657 S/N 3 Mag 2X FAL 79658 S/N 3 Mag 2X

FAL 79659 S/N 4 Mag 2X FAL 79660 S/N 4 Mag 2X

FD 301903

Figure C-4. Grain Etched Condition of Experimental Blades

C-9
ORIGINAL PAGE IS

OF POOR QUALITY

Pratt & Whitney

FR-19009-1

FAL 79689 S/N 5 Mag 2X

FAL 79691 S/N 6 Mag 2X

FAL 79690 S/N 5 Mag 2X

FAL 79692 s/N 6 Mag 2X

FD 301904

Figure C-5. Grain Etched Condition of Experimental Blades

C-10
Pratt & Whitney

FR-19009-1

FAL 79693 S/N 7 Mag 2X FAL 79694 S/N 7 Mag 2X

A B

FAL 79695 S/N 8 Mag 2X

FAL 79697 S/N 8 Mag 2X

FD 301905

Figure C-6. Grain Etched Condition of Experimental Blades

C-ll
Pratt & Whitney

FR-19009-1

FAL 80661 S/N 9 Mag 2X

FAL 80663 S/N 10 Mag 2X

FAL 80662 S/N 9 Mag 2X

FAL 80664 S/N 10 Mag 2X

FD 301906

Figure C-7. Grain Etched Condition of Experimental Blades

C-12
ORIGINAL PAGE IS

OF POOR QUAtrTY

Pratt & Whitney

FR-19009-1

FAL 80665 S/N 11 Mag 2X

FAL 80667 S/N 12 Mag 2X

FAL 80666 S/N 11 Mag 2X

FAL 80668 S/N 12 Mag 2X

FD 301907

Figure C-8. Grain Etched Condition of Experimental Blades

C-13
Pratt & Whitney

FR-19009-1

FAL 80669 S/N 13 Mag 2X FAL 80670 S/N 13 Mag 2X

FD 301908

Figure C-9. Grain Etched Condition of Experimental Blades

C-14