This Project Completion Report was prepared for the National 

Aeronautics and Space Administration, Lewis Research Center. It 

presents the results of a program conducted to establish exothermic 

heated casting technology for the manufacture of low-cost, singlecrystal, uncooled turbine blades for gas turbine engines. The program was conducted as part of the Materials for Advanced Turbine 

Engines (MATE) Program under Contract NAS3-20073, This is Volume 2 

of the Project Completion Report, covering the performance and 

endurance engine testing and post-test component analysis. 

The authors wish to acknowledge the assistance and guidance of 

S. Grisaffe and R. L. Dreshfield of the NASA-Lewis Research Center 

in successfully completing this project. 

PRECEDING PAGE BLA 

iii 

TABLE OF CONTENTS 

Page 

SECTION I 

1.0 SUMMARY 

SECTION I1 

2.0 INTRODUCTION 

SECTION I11 

3.0 FULL-SCALE ENGINE TESTING 

3.1 Performance Testing 

3.2 Endurance Testing 

SECTION IV 

4.0 POST-TEST EVALUATION 

4.1 Interim-Test Blade Evaluation 

4.2 Post-Test Blade Evaluation 

SECTION V 

5.0 CONCLUSIONS 

SECTION VI 

6.0 RECOPllMENDATIONS 

REFERENCES 

1 

5 

7 

10 

17 

23 

23 

21 

35 

37 

39 
LIST OF ILLUSTRATIONS 

Fiqure 

1 

Title 

Cycle A: First 50-Hour Test, High-CycleFatigue Evaluation 

Paqe 

8 

8 

2 Cycle B: Second 50-Hour Test, Stress Rupture 

Evaluation 

Cycle C: Third 50-Hour Test, Simulated 

Commuter Jet Cycle 

3 

9 

Cycle D: Fourth 50-Hour Test, Low-CycleFatigue Evaluation 

4 

9 

5 12 

6 

TFE731 Engine Installed in Phoenix Test Cell 

MATE Rotor Assembly after Performance Testing, 

Showing Tip Rub 14 

Tip View of MATE Rotor Assembly Following 

Performance Test 

7 

15 

HP Turbine Shroud Following Performance Test 

of MATE SC Blades 

8 

16 

9 Sketch Illustrating Flow Discourager Failure 

in Blade N27 During Second 50-Hour Test and 

Sketch Illustrating Flow Discourager Rework 18 

20 

21 

24 

10 

11 

12 

13 

Leading-Edge View of Typical Blades Subjected 

to the Entire 200 Hours of Endurance Testing 

Trailing-Edge View of Typical Blades Subjected 

to the Entire 200 Hours of Endurance Testing 

Visual Examination of SC NASAIR 100 Blade 

(S/N N27) 

Extensive Interconnected and Oxidized Porosity 

and Incipient Melting Were Present in the 

NASAIR 100 Adjacent to the Failed Platform Web 25 

26 

14 Airfoil of Blade S/N N27 Exhibited a 

Microstructure Typical of Fully Processed 

SC NASAIR 100 

Visual Appearance of SC NASAIR 100 (S/N N20, 

N33, N48) and SC Alloy 3 (S/N A42) Blades 

after the 200-Hour Engine Test 

15 

28 

vi 
LIST OF ILLUSTRATIONS (CONTD) 

Fiqure Title Paqe 

16 Microstructure of SC Alloy 3 HP Turbine Blade 

(S/N A42) after 200-Hour Engine Test 29 

17 Microstructure of SC NASAIR 100 HP Turbine 

Blade (S/N N48) after 200-Hour Engine Test 31 

18 Scanning Electron Microscope Photographs 

(2000X) of SC NASAIR 100 HP Turbine Blade 

after 200-Hour Engine Test (S/N N48) 32 

19 Appearance of the Blocky Alpha-Tungsten 

Phase in Airfoil Section of SC NASAIR 100 

HP Turbine Blade 33 

20 SEM Photograph and EDX Analysis of the 

Tungsten + Nickel Rich Mu Secondary Phase 

in the Airfoil Section of the SC NASAIR 100 

Blade (S/N N48) after 200 Hours of Successful 

Engine Testing 34 

vii 
LIST OF TABLES 

Table Title 

1 Blade Substitution Plan for MATE Project 3 

Engine Test 

Paqe 

10 

viii 
SECTION I 

1.0 SUMMARY 

The demand for more efficient and economical engines continues 

to push the industry toward higher turbine operating temperatures. 

The ability to maintain this trend is strongly tied to advancements 

in material technology. Single-crystal (SC) castings offer many 

significant advantages, such as higher melting point and higher 

strength, over: turbine materials currently used in production 

engines. The goals of MATE Project 3 were to develop a low-cost 

casting process capable of producing SC turbine blades and to 

demonstrate the capability of the SC blades through extensive property, rig, and engine testing. 

The following accomplishments are the most significant program 

goals achieved : 

0 Incorporating the MATE high-pressure (HP) turbine blades 

and support hardware into the test engine reduced TSFC by 

1.4 percent and T5 (interturbine temperature) by 2.5 percent, compared to the baseline engine with uncooled 

directlonally solidified (DS) HP turbine blades 

0 The MATE single-crystal turbine blades, both NASAIR 100 

and Alloy 3, successfully completed over 200 hours of 

endurance engine testing with no signs of distress 

o The exothermic casting process was successfully developed into a low-cost nonproprietary method for producing 

sing le-crystal castings 

Project 3 was divided into seven tasks. In Task I, the exothermic casting process was modified to consistently produce 

1 
acceptable so -pressure tur ine blades for the TFE731 

turbofan e ine. This effort focused on the castability and baseline mechanical pro erties of blades produced with MAR-M 247. 

During Task 11, SC alloy derivatives of the MAR-M 247 alloy 

were assessed for castability, microstructure, and improvements in 

mechanical properties. Two of these SC compositions, NASAIR 100 

and Alloy 3, were selected for detailed characterization. 

In Task 111, mechanical, environmental, and physical properties of SC NASAIR 100 and Alloy 3 were quantified. Mechanical 

property testing included creep rupture, tensile, and high- and 

low-cycle fatigue, and the determination of the effects of protective coatings, Environmental’characterization included coated and 

uncoated oxidation and hot-corrosion tests. Physical properties 

measured included density, elastic modulus, thermal expansion, and 

thermal conductivity. 

The SC HP turbine blade was designed in Task IV to utilize the 

improved mechanical properties of the prime SC alloy (NASAIR 100). 

During Task V, two full sets of SC turbine blades and the 

required support hardware for both rig testing and the Task VI 

engine testing were manufactured. Rig and bench testing were also 

completed in Task V. 

The results of the first five project tasks and the achievement of their related goals are presented in Volume 1 of the 

Project 3 Final Report (CR-168218, Garrett No. 21-4314-1). The 

kesults of the final two tasks--Task VI Engine Testing and Task VI1 

Post-Test Evaluation--are presented in this document (Volume 2)- 

The Task VI Engine Testing was successfully completed. Testing consisted of back-to-back performance testing of a production 

2 
configuration TFE731 engine with uncooled DS HP turbine blades and 

the same engine with the ,uncooled MATE SC HP turbine blades and 

related support hardware. This testing showed that the MATE engine 

configuration reduced TSFC by 1.4 percent and T5 by 2.5 percent. 

Performance testing was followed by 200 hours of endurance engine 

testing, divided into the following four 50-hour segments: 

0 50-hours of high-cycle-fatigue (HCF) evaluation 

0 50 

0 50 

0 50 

None of 

ing visually 

was verified 

Task VII. 

hours of maximum power (stress-rupture evaluation) 

hours of simulated commercial mission 

hours of low-cycle-fatigue (LCF) evaluation. 

the blades subjected to the 200 hours of engine testshowed any detrimental effects from the testing. This 

by the post-test metallurgical evaluation performed in 

3 

SECTION I1 

2.0 INTRODUCTION 

The NASA Materials for Advanced Turbine Engines (MATE) Program 

is a cooperative effort with industry to accelerate introduction of 

new materials into aircraft turbine engines, As part of this 

effort, Garrett Turbine Engine Company (GTEG) was authorized under 

NASA Contract NAS3-20073 to develop a new technique for manufacturing low-cost, single-crystal (SC), uncooled cast turbine blades to 

reduce SC casting costs and improve fuel consumption in advanced 

turbofan engines. The process development included those efforts 

required to transfer the technoloqy from the previously demonstrated feasibility stage through component demonstration and 

engine test. Portions of the overall effort included process 

scale-up, alloy evaluations, mechanical property generation, hardware procurement, component testing, and full-scale engine testing 

to evaluate potential benefits. 

This report constitutes Volume 2 of a two-volume Project Completion Report presenting the results of the investigations and 

tests performed under MATE Project 3, Low-Cost Single-Crystal Turbine Blades. This volume covers only the Project 3 full-scale 

engine testing and post-test analysis. All other aspects of this 

project are covered in Volume 1 of this report. 

The intent of Project 3 was to develop a low-cost process to 

produce single-crystal, uncooled turbine blades and to design and 

substitute this blade for the solid, directionally solidified (DS) 

turbine blade used in the high-pressure turbine of the GTEC TFE731 

turbofan engine. 

Project goals associated with this program included the following : 

5 
0 Development of a nonproprietary, lo -cost SC casting process 

0 'Improvement of the stress-rupture capability of the SC 

alloys relative to DS 

0 Design of an uncooled HP turbine blade to use the SC 

material 

0 Demonstration of uncooled SC turbine blades through component and engine testing, 

Project 3 was subdivided into the following seven tasks: 

Task I - Casting Technology 

Task I1 - Alloy/Process Selection 

Task I11 - Property Characterization 

Task IV - Blade Design 

Task V - Component Hanufacture and Testing 

Task VI - Engine Testing 

Task VI1 - Post-Test Evaluation 

Tasks I through V are covered in detail in Volume 1 of this 

Project Completion Report.' In this document, Volume 2, Task VI - 

Full-scale Engine Testing and Task VI1 - Post-Test Evaluation are 

covered in detail, including recommendations concerning the future 

of the exothermic SC casting process and SC HP turbine blades. 

The results of Tasks VI and VII--project completion information--are restricted by the NASA For Early Domestic Dissemination 

(FEDD) policy. The FEDD legend, describing the requirements of 

this policy, is printed on the cover of this document. 

6 
SECTION I11 

3,O FULL-SCALE ENGINE TESTING 

Scope 

The objectives of the Task VI Engine Testing were as follows: 

(1) Verify the anticipated reduction in TSFC with the SC turbine blades 

(2) Demonstrate the durability of both the material and the 

design of the new blade. The program required that the 

HP blades produced in Task V be subjected to 200 hours of 

typical engine operating conditions in an appropriate 

GTEC engine. 

The Task VI Engine Testing of the fully processed HP turbine 

blades consisted of back-to-back performance tests followed by four 

50-hour endurance test segments chosen by GTEC and approved by 

NASA. The performance test was designed to compare a production DS 

turbine configuration with the same engine using the MATE Project 3 

SC turbine blades, The four 50-hour test segments, in the order 

performed, were designed to evaluate the resistance of the SC blade 

to high-cycle fatigue, stress-rupture, a simulated commercial mission, and low-cycle fatigue. These test conditions were chosen to 

allow direct comparison of the SC blades produced in this project 

with the production DS uncooled (solid), MAR-M 247 turbine blades. 

The test cycles for these four test segments are shown in Figures 

1, 2, 3, and 4, The blade substitution schedule presented in Table 

1 was used to expose groups of test blades to single and multiple 

loading conditions established by the test parameters of the four 

50-hour tests. This substitution plan permitted comprehensive 

post-test evaluation of the individual and combined effects of the 

four test segments by both nondestructive and destructive techniques 

7 
13 INCREMENTAL THRUST SETTINGS 

MAX CONTINUOUS 

0 

c 

v) 1.5 HOURS MINIMUM 

oe 

w 

z 

t; 

MINIMUM 

I I I I I 1 

0 10 20 30 40 50 

TIME - HOURS 

Figure 1. Cycle A: First 50-Hour Test, High-Cycle-Fatigue Evaluation. 

MAX CONTINUOUS 

(48 HOURS) 

IDLE 

1 I I I I 

0 10 20 30 40 50 

TIME-HOURS 

Figure 2. Cycle B: Second 50-Hour Test, Stress-Rupture - 

Evaluation. 

8 
CI 

IFn 

Lu 

0 

z 

t; 

a 

3 

n 

IDLE 

ONTl 

7 

MIN 

II 

UOUS 

AX CRUISE 

24 MIN a 

75% CRUISE 

5 5 

I CF 

5 

MIN 

D 

SE 

-0.00 

5 

MIN MIN MIN 

I 

L 

w 

61 50 

I TIME O 

MIN HOURS 

Figure 3. Cycle C: Third 50-Hour Test, Sirnulated Commuter 

Aircraft Mission. 

TOTAL 120 CYCLES 

- 10 MINUTES 

MINIMUM 

MAX CONTINUOUS -- 

10 MINUTES 

MINIMUM 

25 MINUTESICYCLE L 

I 

A 

V 8 I I I 

0 1 20 30 40 50 

TIME - HOURS 

Figure 4. Cycle D: Fourth 50-Hour Test, Low-Cycle-Fatigue 

Evaluation. 

9 
TABLE 1. BLADE SUBSTITUTION PLAN FOR MATE PROJECT 3 ENGINE TEST. 

Material 

NASAIR 100 

Alloy 3 

Totals 

Number of Blades Substituted for Each Test Cycle 

A B C D A+B B+C C+D A+D A+B+C B+C+D A+B+C+D Totals 

222 2 2 14 1 1 18 35 

2 1 1 1 34 39 

2 222*2 2 2* 4* 2 2* 52* 74 

11. I i 

A = 50-Hour High-Cycle-Fatigue Evaluation (Figure 1) 

B = 50-Hour Stress-Rupture Evaluation (Figure 2) 

C = 50-Hour Simulated Mission Evaluation (Figure 3) 

D = 50-Hour Low-Cycle-Fatigue Evaluation (Figure 4) 

*Blades in engine at end of 200-hours 

3.1 Performance Testinq 

The MATE Project 3 SC HP turbine blade is essentially the same 

configuration as the production DS turbine blade, except for the 

addition of a tip winglet and an aft platform flow discourager. To 

accommodate this new blade design, minor modifications were made to 

some of the surrounding hardware. 

10 
To use the superior temperature capability of the SC materials, cooling air supplied to the fore and aft sides of the HP turbine disk was reduced by blocking part of the cooling air entry 

holes. Seals were added between the blades to prevent cooling air 

from leaking out and hot gas from being ingested between the blade 

platforms, and the firtree seal was eliminated to allow more cooling airflow through the firtree. The net effect of these modifications was to reduce the cooling airflow while maintaining adequate 

cooling air to the disk and blade firtree. 

Back-to-back performance tests were conducted to evaluate the 

effect of the SC blade and support hardware on the measured engine 

performance. Testing was done in a fan test cell located in GTEC's 

Phoenix facility, with standard instrumentation. Figure 5 shows 

the TFE731 installed in the fan test cell. 

Following the standard production engine performance baseline 

test, the engine was rebuilt to the MATE Project 3 configuration, 

substituting the following hardware for the original production 

hardware. 

PART NAME 

MATE HPT Blades - NASAIR' 100 

MATE HPT Blades - Alloy 3 

ITT Duct 

HPT Shroud Segments 

LPT Nozzle Duct 

Shroud Retaining Sleeve 

Aft Curvic Coupling 

HPT Seal Plate 

HPT Nozzle, 26 Vane 

11 
Figure 5. TFE731 Engine Installed in Phoenix Test Cell. 

MP-87768 

12 
The engine with this new hardware was then run through the 

standard performance calibration, using the same test cell and 

instrumentation that was used for the baseline performance test. 

Reduction of the performance data to standard-day conditions 

indicated the following: 

0 The baseline 

test limits. 

0 Incorporating 

standard engine was well within production 

the MATE Project 3 SC turbine blades and 

support hardware into the same engine produced a measured 

1.4 percent reduction in TSFC and reduced T5 by 2.5 percent. 

These performance improvements are directly related to the 

increased efficiency of the MATE HP turbine blade and the decrease 

in cooling flow. 

Upon disassembly of the engine, a slight rub was found at 

approximately the 2 o'clock position, forward looking aft. Figures 

6, 7, and 8 show that the assembly and shroud interference was 

minor . 

Visual and fluorescent penetrant inspection (FPI) was completed after teardown, with no distressed parts identified. Several new blades were substituted for endurance testing, and the 

wheel was reassembled and balanced for the first endurance test. 

To avoid further rubs, the ID of the stator shrouds were resized to 

the maximum diameter allowed for production engines. 

13 
Figure 6. MATE Rotor Assembly After Performance Testing. 

m-87767 

14 
MP-07766 

Figure 7. Tip View of MATE Rotor Assembly 

Following Performance Test. 

15 
Figure 8, BP Turbine Shroud Following Performance Test 

MP-87836 of MATE SC Blades. 
3.2 Endurance Testing 

Endurance testing was accomplished in four 50-hour test segments with a teardown inspection after each 50-hour test. Figures 

1, 2, 3, and 4 (previously shown) represent the four test cycles 

used during endurance testing, The blade substitution schedule 

presented in Table '1 was used to expose test blades to various combinations of single and multiple loading conditions to isolate the 

effects of the various tests on the blades. 

All endurance testing was done with standard engine monitoring 

instrumentation at GTEC's remote test facility located at San Tan, 

Arizona. 

The first 50-hour test cycle was established to verify the 

MATE Project 3 blade response and evaluate the resistance to highcycle fatigue (Figure 1). This test was accomplished as scheduled. 

No operating problems occurred during this test, and all test 

parameters were within limits. Post-test inspection revealed no 

distress on any of the SC blades or support hardware. Blades were 

substituted into the rotor assembly in accordance with the previously established substitution plan, the rotor was rebalanced, and 

the engine was reassembled for the second 50-hour endurance test. 

The stress-rupture endurance test cycle used for the second 

test was shown previously in Figure 2, Again, no problems were 

encountered during testing. However, post-test inspection revealed 

that one blade (S/N N27) had lost approximately fifty percent of 

the aft flow discourager. Figure 9A is a sketch of the failed 

blade. No other signs of distress were found on the remaining SC 

blades, but moderate foreign object damage (FOD) was found on the 

first and second LP turbine stages. 

Detailed examination of the failed blade indicated that the 

flow discourager web thickness was undersized and contained excessive porosity due to microshrinkage. Preliminary conclusions were 

that the platform, undersized and weakened by excessive casting 

porosity, could not carry the centrifugal bending load and failed 

17 
at the thinnest section. Further details of this blade examination 

and analysis confirmed the preliminary conclusion and are reported 

in the post-test-evaluation section of this report. 

In addition to the visual and FPI inspections scheduled 

between endurance tests, the web thickness of each blade was measured and the casting porosity in each web was carefully evaluated. 

Three blades were eliminated from further testing due to the combination of a thin platform web and excessive porosity, 

To avoid any chance of additional platform failures, all remaining blades were reworked (as shown in Figure 9B) to reduce the 

Figure 9A. Sketch Illustrating Figure 9B. Sketch Illustrating 

Flaw Discourager Flow Discourager 

Failure in Blade N27 Rework, 

During Second 50-Hour 

Test. 

18 
stress in the flow discourager region. The reworked blades were 

then reassembled into the disk with blades substituted according to 

the schedule. The rotor was rebalanced, the damaged LP turbine 

components replaced, and the engine reassembled for the third 50- 

hour endurance test. 

The third 50-hour test cycle simulated a typical mission for a 

commuter aircraft (Figure 3). No problems were encountered during 

this test, and no significant changes in operating parameters were 

observed as a result of the blade rework. Post-test inspection 

found no signs of distress in the blades or any other component. 

Blade substitutions were again made according to the schedule, the 

wheel rebalanced, and the engine reassembled for the fourth and 

final 50-hour test . 

The fourth 50-hour test cycle, designed to provide an evaluation of the low-cycle-fatigue capability of the SC blades, is shown 

in Figure 4. Normal accelerations and decelerations were maintained for all cycles. No significant engine test problems were 

encountered during this test, and engine performance remained consistent with previous testing, The post-test inspection did not 

reveal any discrepancies in the SC blades or any other hardware. 

Figures 10 and 11 show typical leading and trailing edge views of 

blades tested for the entire 200 hours. 

Detailed examinations of sample test blades were accomplished 

according to the scheduled evaluation. Results of these inspections are presented in the post-test evaluation section of this 

report . 

19 
81674-4 

Figure 10. Leading-Edge View of Typical Blades Subjected to the 

Entire 200 Hours of Endurance Testing. 

MP-87769 

20 
81674-3 

Figure 11. Trailing-Edge View of Typical Blades Subjected 

to the Entire 200 Hours of Endurance Testing. 

MP-87776 

21 

.1 Interim-Test Blade Evaluation 

After 100 hours of testing (Cycles A and B), routine examination of the HP turbine rotor indicated that one of the SC NASAIR 100 

blades (S/N N27) had lost part of the aft flow discourager. The 

weight of the missing piece was estimated at about one gram. To 

identify the cause of this flow discourager failure and to document 

the microstructure of the airfoil, this blade was visually and 

metallurgically examined. These examinations (Figures 12 and 13) 

revealed that the flow discourager web thickness was significantly 

below the minimum specified drawing thickness and only about 65 

percent of the average web thickness. The portion of the platform 

web that failed also had very high amounts of oxidized (surface 

connected) porosity. 

FPI and visual inspection of the remaining blades indicated 

that comparable platform porosity was present in many of the 

blades. However, no other blades had the excessive platform 

porosity and below minimum web thickness that combined to produce a 

failure. 

Microstructural analysis of blade S/N N27 at other locations 

on the blade was also performed (Figures 13 and 14). This analysis 

indicated that the SC NASAIR 100 blade had been solution heattreated at the upper end of the solution heat-treatment window. 

The eutectic gamma prime phase was completely solutioned and minor 

incipient melting was observed in the blade root e Substantially 

less microporosity was noted in the airfoil than in the blade root. 

Energy dispersive X-ray (EDX) analysis was used to verify the 

NASAIR 100 composition of blade S/N N27. 

PRECEDING PAGE BLANK NOT F 

23 

K 
Figure 12. Visual Examination of SC NASAIR 100 Blade (S/N N27) 

Confirms that Substantial Porosity and a Very 

Thin Platform Web Contributed to an Overload 

Fracture of the Flow Discourager Tooth During 

the Second 50-Hour (Maximum Continuous Power) 

Segment of the Engine Test. (Photographs are 

MP-87775 75 percent of indicated magnifications.) 

24 
f 

Figure 13, Extensive Interconnected and Oxidized Porosity (top) 

and Incipient Melting (bottom) Were Present in.the 

NASAIR 10’0 Adjacent to the Failed Platform Web 

(SC NASAIR 100 Blade S/N N27), Photographs are 85 

percent of indicated magnifications, 

25 

MP-90751 
Figure 3.4, Airfoil of Blade S/N N27 Exhibited a 

Microstructure Typical of Fully Processed 

SC NASAIR 100, (Photographs are 80 percent 

of indicated magnifications.) 

Mp-07770 

26 
4-2 Post-Test Blade Evaluation 

Following the removal of the flow discourager tooth edge from 

the remaining SC blades (described previously in this report), the 

last two test cycles (C and D) were completed without incident. At 

the completion of the required 200 hours of testing, all NASAIR 100 

and Alloy 3 turbine blades were carefully examined by both FPI and 

visual inspection at magnifications up to 40X. No evidence of 

distress was noted on any of the blades. 

In addition to the FPI and visual inspections, three SC NASAIR 

100 blades (S/N N20, N33, and N48) and one SC Alloy 3 blade (S/N 

A42), which had been exposed to the full 200 hours of the engine 

test, were selected for metallographic examination. The appearance 

of the blades at the conclusion of the test and the location of the 

longitudinal metallographic section are provided in Figure 15. 

Results of the examination of the SC Alloy 3 blade S/N A42 are 

provided in Figure 16. Metallographic examination indicated that 

the single-crystal alloy was fully solutioned with no evidence of 

incipient melting. Minimal porosity was observed in the airfoil, 

with higher amounts present in the blade root. Scanning electron 

microscopy (SEM) indicated that the cubic gamma prime precipitate 

morphology was retained in the blade root and the minimally 

stressed tip region of the airfoil. In contrast, the cubic gamma 

prime phase transformed to platelets aligned perpendicular to the 

applied centrifugal stress in the [OOl] direction in the higher 

temperature portion of the airfoil. Data in the literature indicates that this gamma prime morphology, commonly known as rafting, 

2,3 occurs at low creep strains and reduces subsequent creep rates. 

Rafting of the gamma prime phase in the airfoil was the only metallurgical modification associated with operation in this engine. 

Post-test metallographic analysis of the SC NASAIR 100 blades 

indicated that they had been solution heat-treated at the upper 

limit of the heat-treatment window. The eutectic gamma prime had 

27 
Figure 15. Visual Appearance of SC NASAIR 100 (S/N N20, N33, 

N48) and SC Alloy 3 (S/N A42) Blades after the 

200-Hour Engine Test. Vertical Lines Indicate 

Location of Longitudinal Metallographic Section. 

MP-87838 

28 
t e 

Mp-90749 

29 
been completely solutioned, and some indications of incipient melting were observed in the three blades examined. Results of the 

microstructural examination of Blade S/N N48 illustrate this condition (Figure 17). In some instances, the amount of incipient 

melting slightly exceeded the one-volume percent permitted by the 

SC NASAIR 100 specification (see Volume 1) .l This condition, however, did not affect the performance of the blades in the engine 

test. 

SEM examination of the SC NASAIR 100 blades indicated that the 

gamma prime phase had undergone modifications similar to the SC 

Alloy 3 material. As shown in Figure 18, the gamma prime phase had 

rafted in the central region of the airfoil. The morphology of the 

rafted gamma prime phase is similar to tht reported by Nathal and 

Ebert3 for NASAIR 100 at the initiation of secondary (steady-state) 

creep. The cubic gamma prime phase morphology, which is typical of 

a fully processed casting, was retained in the relatively cool 

blade root and in the lightly stressed, but hotter, airfoil location near the blade tip. 

Two types of minor secondary phases were also observed in the 

SC NASAIR100 turbine blades. A blocky alpha-tungsten phase, which 

is present in as-cast blades, was observed throughout the root, 

airfoil, and tip (Figures 17 and 18). This phase is shown in a high 

magnification SEM photograph in Figure 19. As expected,. the 

nickel-tungsten-rich Mu phase was also observed in the hotter airfoil sections of the SC NASAIR 100 blades after the 200-hour engine 

exposure (Figure 20). The Mu phase was most commonly observed in 

the airfoil section, where the rafted gamma prime phase was well 

developed (see Figures 18, 20) . 

It should be noted that extensive long-term testing of SC 

NASAIR 100 specimens (see Volume 1) indicated that the Mu phase, in 

the amounts observed, was innocuous.’ Consistent with this information, the Mu phase did not adversely affect the integrity of the 

SC NASAIR 100 KP turbine blades during the 200-hour engine test. 

30 
C A 

Figure 17. Microstructure of SC NASAIR 100 HP Turbine Blade 

(S/N N48) after 200-Hour Engine Test. 

Mp-90750 

31 
MA16568 

Figure 18, Sca icroscope Photographs (2000X) of 

rbine Blade after 200-Hour 

Engine Test (S/N N48), 
Figure 19. Appearance of the Blocky Alpha-Tungsten Phase 

(Arrows) in Airfoil Section of SC NASAIR 100 HP 

Turbine Blade. 

Mp-87765 33 
Figure 20. SEM Photograph and EDX Analysis of the Tungsten+ 

Nickel Rich Mu Secondary Phase in the Airfoil 

Section of an SC NASAIR 100 Blade (S/N N48) 

after 200 Hours of Successful Engine Testing. 

MP-87772 

34 
SECTION V 

5.0 CONCLUSIONS 

The following conclusions were based upon the results of the 

engine testing and post-test analysis of the SC turbine blades presented in this report: 

0 Incorporating the MATE HP turbine blades and support 

hardware into the test engine reduced TSFC by 1.4 percent 

and T5 by 2.5 percent compared to a baseline engine with 

directionally solidified HP turbine blades. 

0 The MATE single-crystal turbine blades, both NASAIR 100 

and Alloy 3, showed no signs of distress after successfully completing over 200 hours of endurance engine 

testing. 

35 

SECTION VI 

6.0 RECOMMENDATIONS 

After completion of the engine testing and the post-test 

evaluation of the SC turbine blades, the following recommendations 

can be made: 

0 Blade Design - The SC blade performed well in the engine 

test, both from efficiency and durability standpoints. 

The only change recommended is a review of the aft flow 

discourager design. Better thickness control or a change 

in thickness is required if the tolerances between the 

cast outer surface or the machined inner surface cannot 

be reduced. 

0 Process Control Plan - Due to the changes in the investment casting industry since this project was initiated, 

it is recommended not to pursue further development of 

the exothermic SC process. The major casting houses have 

already made the capital investment in withdrawal equipment, thus eliminating one advantage of the exothermic 

process. Therefore, unless the proprietary aspect of the 

SC withdrawal process becomes a significant problem, further development of the exothermic process is not recommended. 

0 Property Specifications and Blade Acceptance Criteria - 

No changes are recommended in either of these items at 

this time. However, these areas are constantly reviewed 

and compared to engine test experience. Thus, changes in 

either or both may be desired as the GTEC experience base 

grows. 

PRECEDING PAGE BLANK NOT FILMED 

37 

REFERENCES 

1. Strangman, T.E., B.R. Heath, and M. Fujii: Low-Cost, SingleCrystal Turbine Blades - Volume 1, Garrett Report No, 

21-4314-1, NASA CR-168218, 1983. 

2. Pearson, D.D., F.D. Lemkey and B.H. Kear: "Stress Coarsening 

of Y' and its Influence on Creep Properties of a SingleCrystal Superalloy," in Superalloys 1980 - Proceedings of the 

Fourth International Symposium on Superalloys, Champion, 

Pennsylvania, September 1980, pp:513-520. 

3. NathaP, M.V. and L.J. Ebert, "Gamma Prime Shape Changes During 

Creep of a Nickel-Base Superalloy," Scripta Metallurgica, 17 

(1983), 1151-1154. 

lBUTlON LIST 

NASA Headquarters 

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IBUTlQN LIST (CQNTD) 

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RIBU~I~N LIST (CO 

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Indianapolis, IN 46206 

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Detroit Diesel Allison 

Attn M. Doner 

P.O. Box 894 W!J 

Indianapolis, IN 46206 

45 
DISTRI~UT~ON LIST (CONTD) 

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Indianapolis, IN 46206 

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General Electric Co. 

Attn G. Deboer M8 7 

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General Electric Co. 

Corporate R+D Center 

Attn M. Benz 

P. 0. Box 8 

Schenectady, NY 12301 

Detroit Diesel' Allison 

Attn M. Herman 

P.O. Box 894 

Indianapolis, IN 46206 

General Electric Co. 

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Gas Turbine Prod. aiv. 

1 River Rd. 

Schenectady, NY 12349 

General Electria Co. 

Materials and PrOCeSSihg Attn R. Sprague M82 

Evandale, OH 45215 

General Electria Coahpen Attn Tech Info Cntt H-!Z 

Evandale, OH 45215 

General Electric Co. 

Attn W. E. Kurz 

Mgr. MDO 

1000 Western Ave. 

Lynn, MR 01910 

General Electric Co. 

Corporate RtD Centet 

Attn M.F.X. Gigliotti P. 0. Box 8 

Schenectady, NY 12381 

Pratt and Whitney Aircraft Pratt and Whitney Aircraft 

Commercial Products Commercial Products 

Attn A. Hauser Attn M. Gel1 

400 Main St. 400 Main St. 

East Hartford, CT 06108 East Hartford, CT 06168 

Detroit Diesel Allison Detroit Diesel Allison 

Attn N. Provenzano U29A Attn J. Lyon 

P.O, Bow a94 P.O. Box 894 

Indianapolis, IN 46206 Indianapolis, IN 96206 

46 
DISTRIBUTION LlST (CONTDI 

General Electric Co. 

Attn F. D. Lordi, Bldg 53 

Gas Turb Prod. Div. 

SchenectGdp, NY 12345 

General Electric Co. 

Material and Processing 

Technology 

Attn T, F. Berry 

Evandale, OH 45215 

General Electric Co. 

Material and Processing 

Technolo 3 Laboratory 

Attn Nc fairbanks 

Evandale, OH 45215 

,I 

General Electrio Co. 

Attn J. Hsla 

AEG/GE1) 

1000 Western Ave. 

Lynn, MA 01910 

General Eledtria Cob 

Corporate R+D Center 

Attn Libraty 

P. 0. BOH 8 

Schenectady. NY 12301 

Pratt and Whitney Aircraft 

Commercial Products 

Attn C. P. Sullivan 

400 Main St. 

East Hartford, CT 06108 

Pratt and Whitney Aircraft 

Manufacturing DLV. 

Attn F. M. Heinrich 

400 Main St. 

East Hartford, CT 06040 

Pratt and Whitne Aircraft 

Attn H. Prziredel 

Box 2691 

West Palm Beach 

FL 33402 

General Electric Co. 

Attn W. Schilling 

Bldg 53 

Gas Turb Prod. Div. 

Schenectady, MY 12345 

General Electric Co. 

Material and Processing Technolog Laboratory 

Attn L. Wflbers M87 

Evandale, OH 45215 

General Electritt eo. 

Material and Processing Technology Laboratory 

Attn R. E. Allen 

Evandale, OH 45215 

General Electrio Co. 

Attn 0. Steele Irons 

Mgr. Casting Tech. 

1000 Western Ave. 

Lynn, MA 01910 

Pratt and Whitney Aircraft Commercial Ptodu6ts 

Attn S. Blecherman 

Aircraft Road Bldg 140 Middletown 

CT 06457 

Pratt and Whitne Aircraft 

Manufacturing D Y v. 

Attn J. Zubeckis 

400 Main St. 

East Hartford, CT 06108 

Pratt and Whitney Aircraft 

Attn J. Moore 

Box 2691 

West Palm Beach 

FL 33402 

Solar 

Attn A. Stetson 

P.O. Box 80966 

San Diego, CA 92138 

47 
DISTR~BUTION LIST (CONTD) 

Williams Research 

Attn Wm. P. Schimmel 

2280 W. Maple Walled Lake, HI 48088 

Williams Research 

Attn J. Carlton 

2280 W. Maple Walled Lake, HI 48088 

Battelle Memorial Inst 

Attn D. Niesz 

505 King Ave. 

Columbus, OH 43201 

Williams Research 

Attn Librar 

2280 W. HapYe Walled Lake, nT SbbUU 

Airco Temescal 

Attd L. Bianchi 

2850 Seventh St. 

Berkeley, CA 94710 

Battelle Memorial Inlit 

Attn Technical LibrQry 

505 King Ave. 

Columbus, OH 43261 

The Boeing Co Boeing - Wichita biv, Attn H. D. Kevin 41-SH Attn Wm. F. Schmidt 

P.O. Box 3999 org. 75930 Propulsion Seattle, WA 98124 Wichita, KS 67210 

Cameron Iron Works 

Attn Library 

P. 0. Box 1212 

Houston, TX 77001 

Cannon-Muskegon Cotp. Attn 6. L. Erickson 

P.O. Box 506 

Huskegon, HI 49443 

Pratt.and Whitney Aircraft Pratt and Whitney Aircraft 

Attn H. 1. Hess, Jr Attn J. Winfree 

Box 2691 Box 2691 

West Palm Beach West Palm Beach 

FL 33402 fL 33402 

Solar Teledyne CAE 

PO Box 6971 

Toledo, OH 43612 

Attn A. Metcalfe Hat. DeV. and Msnuf. Engr P.O. Box 80966 Attn R. Beck 

San DiegB,’ CA 92138 

Williams Research Williams Research 

Attn 4. Jones Attn P. Nag 2280 W. Maple Walled %take, MI 48088 Walled fake, kC 48088 

2280 W. Mapfe 

48 
DISTRIBUTION LiST (CONTD) 

Airco %'ernescal 

Attn W.R. Halnan 

2850 Seventh 3t. 

Berkeley# CA 94710 

Battelle Memorial hst 

Attn K. Meiners 

505 King Ave. 

Columbus, OH 43201 

BoeingiComm. Air lane Co P Beech Aircraft Cor oration 

Attn C A, Remblesle Attn W. Blissel 9H-43 

9709 E. Central Ave. P.O. Box 3767 

Wichita, KS 67201 Seattler WA 98124 

Brown Boveri Turbo. Inc. Brunswick Carp. 

Attn A. Giamarilse Attn. R. Tolokan 

711 Anderson Ave. N Dehndr FL 32720 

L. Engel 2000 Brunswick Lane 

St. Cloud, 56301 

Cannon-Mugkegon Corp. Carpenter Technology Corp. 

Attn R. Quigg Attn 0. DelCorso 

P.O. BOW SO6 A. Walsh 

Muskegonr M 49443 RID Labr Bldg. 68 

Readlng, PA 19603 

Cartech Inc. 

Attn E. Berger 

P.O. Box 662 

Reading, PA 19603 

Chrysler Corp. 

Attn F.A. Hagen 

BOX 1118 CIMS 418-19-30 

Detroit, MI 48288 

DCI Inc. 

Attn R. Engdahl 

318 Victory Drrve 

Herndon, VA 22070 

Duradyne Tech. Inc. 

Attn R. Horton 

8607 Tyler Blvd 

Mentor 

OH 44060 

Cessna Aircraft Co. 

Attn A. Kavie 

(Turbofan) 

Wallace Division 

P.O. Box 7704 

Wichita, KS 67277 

Climax Molybdenum Co. 

Attn Wm. Hagel 

P.O. Box 1568 

Ann Arbor, HI 48106 

Defense Marketing Services 

Attn D. Franus 

100 Northfield St. 

Greenwich, CT 06830 

Engelhard Industries 

Attn E. Grider 

Route 152 

Plainville, HA 02762 

49 
DISTRIBUT~ON LIST (CONTD) 

Gates Learjet Corp. 

Attn W. J. Reese 

P.O. Box 7707 

Wichita KS 67277 

C. L. King 

General Motors Corp. 

Attn A. Bell 

Advanced Prod Engr Technical Center 

Warren, HI 48090 

Grumman Aerospace Corp. 

Attn C. Hoelzer C32-5 

Bethpage, NY 11719 

Cessna Airctaft Co. 

Attn C. Gonzales 

(Turboprop) 

Wallace Division 

P.O. Box 7704 

Wichitar KS 67277 

Gates Learjet Corp. 

Attn T. Reichenberget 

P.O. Box 7707 

Wichita, KS 67277 

General Motors Corp. 

attn K. Bly 73-5 

Technical Center 

Warren, MI 48090 

Howmet Turbine Comp. biv. 

Attn W. Freeman 

475 Steamboat Road 

Greenwich, CT 06830 

Chromoloy 

Attn D. Kenton 

P,O. Box 10699 

Midwest City, OK 73140 

Colt Industries Colt Industries 

Attn E. Dulis Attn 3. Moll 

Crucible-Inc, Mat'ls Rsch 

P. 0. Box 88 P: 0. Box 88 

Pittsburgh, PA 15230 Pittsburgh, PA 15230 

Crucible Inc, Mat'ls Rsch 

Delta Airlines 

Attn W.J. Overend 

Atlanta Airport 

Atlanta, 43A 30320 

Ford Mota? Company 

Attn C. Feltner 

Metallurgical Dept. 

P.O. Bo% 2053 

Dearborn, Mf 98121 

Douglas Aircraft Co. 

Attn Library 

McDonnell Douglas Corp 

3855 Lakewood Blvd. 

Long Beach, CA 90846 

Ford Motor Company 

Attn Y. Telang Metallurgical Dept. 

P.O. Box 2053 

Dearborn, MI 48121 

Gates Learjet Corp. General Atomics Co. 

Attn A. M. Heinyich Attn D. I. Robert& 

T. Reichenberger P. 0. Box 81608 

P.O. Box 7707 San Diego, CA 92138 

Wichita, US 67217. 

50 
D~ST~~BUTIQN LIST (CQNTD) 

Gould In@. , Attn E. N. A ua 

540 E. 105 Sf 

Cleveland, OH-49106 

Howmet Corporation Technical Center 

,Attn L. Dardi 

699 Benston Rd. 

Whitehall, MI 49461 

Howmet Corporation Turbine Components Div. 

Attn Mr. Burd 

Dover, NJ 07801 

Huntington Alloys Div. 

International Nickel Co. 

Attn J. Cundiff 

Huntington, WV 25720 

Intermetco 

Attn J. Siergiej 

300 Concord Road 

Wayland, MA 01778 

Kelsey Hayes Co. 

Attn Wm. E. Good 

Heintz Div. 

Front St. and Olney he. 

Philadelphia, PA 19120 

METCO Inc. 

Attn V. Lanza 

1101 Prospect Ave. 

Westbury, L.I., NY 11590 

Rand Corp. 

Attn J. Richard Nelson 

Washington Research Div. 

2100 M St. 

Washington, DC 20037 

Gould Inc. 

Attn G. Davies 

540 E. 105th St. 

Cleveland, OH 44108 

Howmet Corporation Technical Center 

Attn J. VanderSluis 

699 Benston Rd. 

Whitehall, MI 49961 

Howmet Corporatidn Turbine Components Div. 

Attn E. Carozza 

Mr. Burd 

Dover, NJ 07801 

IIT Research Institute 

Attn M. A. Howes 

10 West 35th Street 

Chicago, IL 60616 

jetshapes, Inc. 

Attn F. Jaeger Rockleigh Industrial Perk 

Rockleigh, NJ 07647 

Thomas McNamara 

Consultant 

42701 Main St.* SP 154 

Sen Jacinto 

CA 92383 

METCO Inc. 

Attn C. Lewis 

3400 A Oak Cliff Rd. 

Atlanta, GA 30340 

Rocketdyne Division 

Attn J. Frandsen 

Rockwell International 

6633 Canoga Ave. 

Canoga Park, CA 91364 

51 
5lSTRlBUTlON LIST (CONTD) 

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Attn L.L. McHughes General Aviation Div. 

5001 N. Rockwell Ave. 

Bethany, OK 73008 

Huntington A110 s Div. 

International Kickel Co. 

Attn R. Haeberle, 3r 

Huntington, WV 25720 

IIT Research Institute 

Attn V. Hill 

10 West 39th Street 

Chicago, IL 60616 

Kaman Sciehces Corp. 

Attn D. H. Seita 

P. 0. Bow 7463 

Colorado Springs 

60 80933 

MCIC 

Battelle Memorial Inst. 

Attn H. Mihdlin 

COlUmbUSr Ofl 93201 

Pacifia Southwest Airlines 

Attn E. Chambers 

7707 Consolidated Way 

San Diego, CR 92121 

Rockwell International 

Attn V. Ta lor 

General Avration Div. ' 

5001 N. Rockwell Ave. 

Bethany, OK 73008 

Cabot Cor 

Attn D. tiaistrom 

1020 W, Park Xve. 

Kokomar IN 46901 

Special Metals, Stla, 

Attn J. Pridgeon Middle Settlement Road 

New Hartford 

Neu York 13413 

Huntington Allo s Div. 

International Kickel Co. 

Attn F. Perry Huntington, WV 25720 

IIT Research Institute 

Attn. R. Firestone 

10 West 35th St. 

Chicago 

IL 60616 

Kelsey Hayes Go. 

Attn W.G, Koby Heinta Div. 

Front St. and Olney Ave. 

Philadelphia, PA 19120 

Martin Marietta 

Attn C.H. Lund 

15 N. Windsor Rd. 

Arlington Hts 

IL 60004 

Pan American Woi.16 lirways 

Attn W.B. Hibbs 

Pan Am Buildin 

New York. NY fool7 

Rockwell International 

Attn. E. Mathwi 

General Rviation #iv . 5001 N. Rockwell Ave. 

Bethany, OK 73008 

Stellite Division 

Cabot Corporation 

Attn M. Rothman 

1020 West Park he. 

Kokomo, IN 46901 

52 
~~STRIBUT~ON LIST (CONTD) 

Sundstrand 

Attn D. Augustine 

4747 Harrison Ave 

Rockford, IL 61101 

TRW Inc . Attn C.S. Kortovich 

23555 Euclid Ave. 

Cleveland, OH 44117 

TRW Inc . Attn T. Piwonka 

23555 Euclid Ave. 

Cleveland, OH 44117 

Turbine Support Co. 

Attn R. Van Cleaf 

P. 0. Box 20148 

P.O. Box 20148 

San Antonio, TX 78220 

United Tech Rsch Center 

Attn B. Thomson 

East Hartford, CT 06108 

Universal Cyclops 

Attn R. Gasior 

Mayer Street 

Bridgville, PA 15017 

Vought Corp. 

Attn Library 

2-50370/TL 7-67 

P.O. Box 5907 

Dallas TX 75222 

Teledyne Allvaa 

Attn S. C. Pearman 

P. 0. Box 759 

Monroe, NC 28110 

Swearingen Aviation Cotp. 

Attn R. McKelvey 

P.O. Box 32486 

San Antonio, TX 78284 

TRW Inc . Attn J. McCarthy 

23555 Euclid Ave. 

Cleveland, OH 4411) 

TRW Inc . Attn C. Vishnevskv 

23555 Euclid Ave.- 

Cleveland, OH 44ll# 

Union Carbide Corp. eLinde Division 

441 Sackett Pt. Rd. 

North Haven, CT 06475 

Attn: E. B. Cook, Jt. 

United Tech Rsch Center 

Attn Library 

East Hartford, Ct bSlOb 

Universal Cyclops 

Attn Wm. Kent 

Player St 

Bridgeville 

PA 15017 

Westinghouse R + t, Centei: 

Attn D. Moon 

Beulah Rd. 

Pittsburgh, PA 15235 

Teledyne Allvac 

Attn R. Kennedy 

Wm. Thomas 

P. 0. Box 759 

Monroe, NC 28110 

53 
DIST~IBUTIQN LIST (CQNTD) 

TRW Inc . TRW Inca 

Attn C. Berth Attn D. Maxwell 

23555 Euclid Ave. 23555 Euclid Ave. 

Cleveland, ' OH 44 11 7 Cleveland, OH 44117 

TRW In&. TRW Inc . Attn E. A. Steigerwald Attn J. Preston 

Metals Oivision Metals Division 

Minerve, OH 94657 Minerva, OH 44657 

United Airline United Airlines - SFOEG 

Attn Y. Tekerian Attn J. IC. Goodwine 

San.Frencisc!o Airport Sen Francisco Airport 

Sen Francisco CA 94128 

CA 99128 

United tech Lsch Center Universal Cyclops 

Attn F. D. Lemkey Attn L. therbier 

East Hartford, CT 06108 Player St. 

Bridgeville 

PA 15017 

Vought Carp. Vought Corp. 

Attn W.R. Boruff 2-53220 Attn.0.H. Cook 2-53400 

P.O. BOW 3907 P.O. Box 5909 

Dallas TX 75222 Dallas TX 95222 

Westingheuse Electric Co. Westinghouse Eleotric Co. 

Attn R. L. Ammon Attn D. Goldberg 

P,O. Bolt 10864 

Pxttsbutgh, PA 19236 Pittsburgh, PA 15236 

P,O, Box l086Q 

Westinghouse Electric Co. , Attn E. Crombie C-210 

P.O. Box 251 

Concordville, PA 19331 

54 
1065 

Unclassified 

1. Repart No 

CR-17 4 6 52 

2. Government Accession No. 3. Recipient's Catalog No. 

Unclassified 1 61 

4. Title and Subtitle 

Low-Cost S ing le-Cry stal Turbine Blades 

llolume 2 

April 1984 

6. Performing Organization Code 

7. Author(s1 8. Performing Organization Report No. 

T. E. Strangman R. E. Dennis B. R. Heath Garrett 21-4314-2 

. 10. Work Unit No. 

9. Performing Organization Name and Address 

Garrett Turbine Engine Company 

A Division of The Garrett Corporation 

Phoenix, Arizona 85010 NAS3-20073 

12. Sponsoring Agency Name and Address 

Washington, D.C. 20546 

11. Contract or Grant No 

13. Type of Repon and Period Covered 

Project Completion Report 

Project 3 

National Aeronautics and Space Administration 14. Sponsoring Agency Code 

15. Supplementary Notes 

Project Manager: Robert L. Dreshfield, Materials Division 

NASA-Lewis Research Center, Cleveland, Ohio 

16. Abstract 

The overall objectives of Project 3 were to develop the exothermic casting process to produce uncooled single-crystal (SC) HP turbine blades in MAR-M 247 and higher 

strength derivative alloys and to validate the materials process and components 

through extensive mechanical property testing, rig testing, and 200 hours of endurance engine test ing. 

These Program objectives were achieved. The exothermic casting process was 

successfully developed into a low-cost nonproprietary method for producing singlecrystal castings. 

Single-crystal MAR-M 247 and two derivative DS alloys developed during this 

project, NASAIR 100 and SC Alloy 3, were fully characterized through mechanical 

property testing. SC MAR-M 247 shows no significant improvement in strength over 

directionally solidified (DS) MAR-M 247, but the derivative alloys, NASAIR 100 and 

Alloy 3, show significant tensile and fatigue improvements. 

Firtree testing, holography, and strain-gauge rig testing were used to determine 

the effects of the anisotropic characteristics of single-crystal materials. Na 

undesirable characteristics were found. In general, the single-crystal material 

behaved similarly to DS MAR-M 247. Two complete engine sets of SC HP turbine blades 

were cast using the exothermic casting process and fully machined. These blades were 

successfully engine-tested. 

17. Key Words (Suggested by Author($) I 

Tur b ine-Blade 

S ing le-Cr y s t a1 

Alloy-Selection NASAIR 100 

Ni-base Alloys Alloy 3 

Investment Casting 

Exothermic-Heatinq 

19. Security Classif. (of this report) 

18. Distribution Statement 

Columnar -Gr a in 

MAR-M 247 

I 20. Security Classif. (of this page) 1 21. NO. of Pages I 22. Rice' 

NASA-(2-168 (Rev. 10-75)