Low-cost single-crystal turbine blades, volume 2

TABLE OF CONTENTS Page SECTION I 1 . 0 SUMMARY SECTION I1 2 . 0 INTRODUCTION SECTION I11 3.0 FULL-SCALE ENGINE TESTING 1 0 3.1 P e r f o r m a n c e T e s t i n g 3.2 E n d u r a n c e T e s t i n g SECTION I V 4.0 POST-TEST EVALUATION 4.1 I n t e r i m - T e s t B l a d e E v a l u a t i o n 4.2 P o s t - T e s t B l a d e E v a l u a t i o n SECTION V 5.0 CONCLUSIONS SECTION V I 6.0 RECOPllMENDATIONS REFERENCES LIST OF ILLUSTRATIONS Title Paqe Fiqure 1 Cycle A : First 50-Hour Test, High-Cycle- Fatigue Evaluation Cycle B: Second 50-Hour Test, Stress Rupture Evaluation Cycle C : Third 50-Hour Test, Simulated 3 Commuter Jet Cycle Cycle D: Fourth 50-Hour Test, Low-Cycle- Fatigue Evaluation 5 TFE731 Engine Installed in Phoenix Test Cell MATE Rotor Assembly after Performance Testing, Showing Tip Rub Tip View of MATE Rotor Assembly Following Performance Test HP Turbine Shroud Following Performance Test of MATE SC Blades Sketch Illustrating Flow Discourager Failure in Blade N27 During Second 50-Hour Test and Sketch Illustrating Flow Discourager Rework 18 10 Leading-Edge View of Typical Blades Subjected to the Entire 200 Hours of Endurance Testing 11 Trailing-Edge View of Typical Blades Subjected 21 to the Entire 200 Hours of Endurance Testing 12 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 14 Airfoil of Blade S/N N27 Exhibited a Microstructure Typical of Fully Processed 26 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 vi LIST OF ILLUSTRATIONS (CONTD) F i q u r e T i t l e P a q e 1 6 Microstructure o f SC A l l o y 3 HP T u r b i n e Blade (S/N A42) a f t e r 200-Hour E n g i n e T e s t 29 17 Microstructure o f SC NASAIR 1 0 0 HP T u r b i n e Blade ( S / N N 4 8 ) after 200-Hour E n g i n e T e s t 31 18 S c a n n i n g E l e c t r o n Microscope P h o t o g r a p h s (2000X) o f SC NASAIR 1 0 0 HP T u r b i n e Blade a f t e r 200-Hour E n g i n e T e s t (S/N N48) 3 2 1 9 A p p e a r a n c e o f t h e B l o c k y Alpha-Tungsten P h a s e i n A i r f o i l S e c t i o n o f SC NASAIR 1 0 0 HP T u r b i n e B l a d e 33 20 SEM P h o t o g r a p h a n d EDX A n a l y s i s o f t h e T u n g s t e n + N i c k e l R i c h M u S e c o n d a r y P h a s e i n t h e A i r f o i l S e c t i o n o f t h e SC NASAIR 1 0 0 B l a d e ( S / N N48) a f t e r 200 H o u r s o f S u c c e s s f u l E n g i n e T e s t i n g v i i LIST OF TABLES Table Title Paqe 1 Blade Substitution Plan for MATE Project 3 Engine Test viii

SECTION I 1.0 SUMMARY The demand for more e f f i c i e n t and economical e n g i n e s c o n t i n u e s to push t h e i n d u s t r y toward h i g h e r t u r b i n e o p e r a t i n g t e m p e r a t u r e s .

The a b i l i t y to m a i n t a i n t h i s t r e n d is s t r o n g l y t i e d to advancements i n material technology. S i n g l e - c r y s t a l (SC) c a s t i n g s o f f e r many s i g n i f i c a n t a d v a n t a g e s , such as h i g h e r m e l t i n g p o i n t and h i g h e r s t r e n g t h , over: t u r b i n e materials c u r r e n t l y used i n p r o d u c t i o n e n g i n e s . The goals of MATE Project 3 were to d e v e l o p a low-cost c a s t i n g process c a p a b l e of producing SC t u r b i n e blades and t o d e m o n s t r a t e t h e c a p a b i l i t y of t h e SC blades t h r o u g h e x t e n s i v e prop- e r t y , r i g , and e n g i n e t e s t i n g .

The f o l l o w i n g accomplishments a r e t h e most s i g n i f i c a n t program g o a l s a c h i e v e d : 0 I n c o r p o r a t i n g t h e MATE h i g h - p r e s s u r e ( H P ) t u r b i n e blades and s u p p o r t hardware i n t o t h e t e s t e n g i n e reduced TSFC by 1 . 4 p e r c e n t and T5 ( i n t e r t u r b i n e t e m p e r a t u r e ) by 2.5 per- c e n t , compared to t h e b a s e l i n e e n g i n e w i t h uncooled d i r e c t l o n a l l y s o l i d i f i e d ( D S ) H P t u r b i n e blades 0 The MATE s i n g l e - c r y s t a l t u r b i n e blades, both NASAIR 1 0 0 and A l l o y 3 , s u c c e s s f u l l y completed o v e r 200 h o u r s of e n d u r a n c e e n g i n e t e s t i n g w i t h no s i g n s of d i s t r e s s o The e x o t h e r m i c c a s t i n g process was s u c c e s s f u l l y d e v e l - oped i n t o a low-cost n o n p r o p r i e t a r y method for p r o d u c i n g s i n g l e - c r y s t a l c a s t i n g s Project 3 was d i v i d e d i n t o s e v e n t a s k s . I n T a s k I, t h e exo- t h e r m i c c a s t i n g process w a s modified to c o n s i s t e n t l y produce acceptable so -pressure tur ine blades for the TFE731 turbofan e ine. This effort focused on the castability and base- line 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 proper- ties 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 protec- tive 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 achieve- ment 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. Test- ing consisted of back-to-back performance testing of a production 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 hours of maximum power (stress-rupture evaluation) 0 50 0 50 hours of simulated commercial mission 0 50 hours of low-cycle-fatigue (LCF) evaluation.

None of the blades subjected to the 200 hours of engine test- ing visually showed any detrimental effects from the testing. This was verified by the post-test metallurgical evaluation performed in Task VII.

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 manufactur- ing 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 demon- strated feasibility stage through component demonstration and engine test. Portions of the overall effort included process scale-up, alloy evaluations, mechanical property generation, hard- ware procurement, component testing, and full-scale engine testing to evaluate potential benefits.

This report constitutes Volume 2 of a two-volume Project Com- pletion Report presenting the results of the investigations and tests performed under MATE Project 3, Low-Cost Single-Crystal Tur- bine 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 fol- lowing : 0 Development of a nonproprietary, lo -cost SC casting pro- cess 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 com- ponent and engine testing, Project 3 was subdivided into the following seven tasks:

- Casting Technology

Task I

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 informa- tion--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.

SECTION I11 3,O FULL-SCALE ENGINE TESTING Scope The objectives of the Task VI Engine Testing were as follows: Verify the anticipated reduction in TSFC with the SC tur- (1) bine 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 mis- sion, 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 tech- niques 13 INCREMENTAL THRUST SETTINGS MAX CONTINUOUS

z

c t; v) 1.5 HOURS MINIMUM oe w MINIMUM I I I I I 0 10 20 30 40 50 TIME - HOURS F i g u r e 1 . C y c l e A: F i r s t 50-Hour T e s t , High-Cycle-Fatigue E v a l u a t i o n .

MAX CONTINUOUS (48 HOURS) IDLE I I I I 0 10 20 30 40 50 TIME-HOURS F i g u r e 2. C y c l e B: Second 50-Hour T e s t , S t r e s s - R u p t u r e - E v a l u a t i o n .

ONTl UOUS II AX CRUISE MIN CI 24 MIN a I CF SE z D I- 75% CRUISE t; Fn MIN a Lu n -0.00 IDLE 5 5 5 MIN MIN MIN L w

I I

TIME 61 50 O MIN HOURS F i g u r e 3 . C y c l e C: T h i r d 50-Hour T e s t , Sirnulated C o m m u t e r A i r c r a f t Mission.

TOTAL 120 CYCLES

- 10 MINUTES

MINIMUM MAX CONTINUOUS

- -

10 MINUTES MINIMUM L 25 MINUTESICYCLE A I 8 I I I V 0 1 2 0 30 40 50

TIME - HOURS

F i g u r e 4 . C y c l e D: F o u r t h 50-Hour T e s t , Low-Cycle-Fatigue E v a l u a t i o n .

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

Number of B l a d e s S u b s t i t u t e d for E a c h T e s t C y c l e A B C D A+B B+C C+D A+D A+B+C B+C+D A+B+C+D T o t a l s Material NASAIR 100 2 2 2 2 2 1 4 1 1 18 35 A l l o y 3 2 1 1 1 34 39 T o t a l s 2 2 2 2 * 2 2 2* 4* 2 2* 52* 74 1 1 . I i A = 50-Hour H i g h - C y c l e - F a t i g u e E v a l u a t i o n ( F i g u r e 1) B = 50-Hour Stress-Rupture E v a l u a t i o n ( F i g u r e 2 ) C = 50-Hour Simulated M i s s i o n E v a l u a t i o n ( F i g u r e 3) D = 50-Hour Low-Cycle-Fatigue E v a l u a t i o n ( F i g u r e 4 ) * B l a d e s i n e n g i n e a t e n d of 200-hours 3.1 P e r f o r m a n c e T e s t i n q The MATE P r o j e c t 3 SC H P t u r b i n e b l a d e is e s s e n t i a l l y t h e same c o n f i g u r a t i o n as t h e p r o d u c t i o n D S t u r b i n e b l a d e , except f o r t h e of a t i p w i n g l e t and a n a f t p l a t f o r m f l o w d i s c o u r a g e r . T o a d d i t i o n accommodate t h i s new blade d e s i g n , minor modifications were made t o some o f t h e surrounding h a r d w a r e .

T o u s e t h e s u p e r i o r t e m p e r a t u r e c a p a b i l i t y of t h e SC materi- a l s , c o o l i n g a i r s u p p l i e d to t h e fore and a f t sides of t h e H P t u r - b i n e d i s k was reduced by b l o c k i n g part of t h e c o o l i n g air e n t r y holes. Seals were added between t h e blades to p r e v e n t c o o l i n g a i r from l e a k i n g o u t and h o t g a s from b e i n g i n g e s t e d between t h e blade platforms, and t h e f i r t r e e seal was e l i m i n a t e d to allow more cool- i n g airflow t h r o u g h the firtree. The n e t effect of these m o d i f i c a - t i o n s was to r e d u c e t h e c o o l i n g airflow w h i l e m a i n t a i n i n g a d e q u a t e c o o l i n g a i r to t h e d i s k and blade firtree.

Back-to-back performance tests were conducted t o e v a l u a t e t h e e f f e c t of t h e SC blade and s u p p o r t hardware on t h e measured e n g i n e performance. T e s t i n g w a s done i n a f a n test cell located i n GTEC's Phoenix f a c i l i t y , w i t h s t a n d a r d i n s t r u m e n t a t i o n . F i g u r e 5 shows t h e TFE731 i n s t a l l e d i n t h e f a n test cell.

Following t h e s t a n d a r d p r o d u c t i o n e n g i n e performance b a s e l i n e test, t h e e n g i n e was r e b u i l t to t h e MATE Project 3 c o n f i g u r a t i o n , s u b s t i t u t i n g t h e f o l l o w i n g hardware f o r t h e o r i g i n a l p r o d u c t i o n hardware.

PART NAME

MATE HPT Blades - NASAIR' 100

MATE H P T Blades - A l l o y 3

ITT Duct HPT Shroud Segments LPT N o z z l e Duct Shroud R e t a i n i n g S l e e v e A f t C u r v i c Coupling H P T Seal Plate HPT N o z z l e , 26 Vane Figure 5 . TFE731 Engine I n s t a l l e d i n Phoenix T e s t C e l l .

MP-87768 The e n g i n e w i t h t h i s new hardware w a s t h e n r u n t h r o u g h t h e s t a n d a r d performance c a l i b r a t i o n , u s i n g t h e same test cell and t h a t was u s e d for t h e b a s e l i n e performance test.

i n s t r u m e n t a t i o n R e d u c t i o n o f t h e performance data to standard-day c o n d i t i o n s i n d i c a t e d t h e f o l l o w i n g : 0 The b a s e l i n e s t a n d a r d e n g i n e w a s w e l l w i t h i n p r o d u c t i o n test l i m i t s .

0 I n c o r p o r a t i n g t h e MATE Project 3 SC t u r b i n e blades and s u p p o r t hardware i n t o t h e same e n g i n e produced a measured 1.4 p e r c e n t r e d u c t i o n i n TSFC and reduced T5 by 2.5 per- c e n t .

These performance improvements are d i r e c t l y related to t h e i n c r e a s e d e f f i c i e n c y of t h e MATE HP t u r b i n e blade and t h e decrease i n c o o l i n g flow.

Upon d i s a s s e m b l y of t h e e n g i n e , a s l i g h t r u b was found a t a p p r o x i m a t e l y t h e 2 o'clock p o s i t i o n , forward l o o k i n g a f t . F i g u r e s 6 , 7 , and 8 show t h a t the assembly and shroud i n t e r f e r e n c e was minor .

V i s u a l and f l u o r e s c e n t p e n e t r a n t i n s p e c t i o n ( F P I ) was com- pleted after teardown, w i t h no distressed p a r t s i d e n t i f i e d . Sev- e r a l new blades were s u b s t i t u t e d for e n d u r a n c e t e s t i n g , and t h e wheel w a s reassembled and b a l a n c e d f o r t h e f i r s t endurance test.

T o a v o i d f u r t h e r r u b s , t h e I D of t h e stator s h r o u d s were r e s i z e d to t h e maximum diameter allowed for p r o d u c t i o n e n g i n e s .

MATE Rotor Assembly After Performance T e s t i n g .

F i g u r e 6 .

m-87767 F i g u r e 7. T i p V i e w of MATE Rotor Assembly MP-07766 Following P e r f o r m a n c e T e s t .

F i g u r e 8 , BP Turbine Shroud Following Performance T e s t MP-87836 of MATE SC Blades.

3 . 2 Endurance Testing Endurance testing was accomplished in four 50-hour test seg- ments 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 com- binations 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 high- cycle fatigue (Figure 1). This test was accomplished as scheduled.

N o 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 previ- ously 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. N o 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 exces- sive 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 a t t h e t h i n n e s t s e c t i o n . F u r t h e r d e t a i l s o f t h i s blade e x a m i n a t i o n and a n a l y s i s confirmed t h e p r e l i m i n a r y c o n c l u s i o n and are reported i n t h e p o s t - t e s t - e v a l u a t i o n s e c t i o n of t h i s report.

I n a d d i t i o n to t h e v i s u a l and FPI i n s p e c t i o n s s c h e d u l e d between endurance tests, t h e web t h i c k n e s s of each blade was mea- s u r e d and t h e c a s t i n g porosity i n each web was c a r e f u l l y e v a l u a t e d .

Three blades were e l i m i n a t e d from f u r t h e r t e s t i n g due to t h e com- b i n a t i o n of a t h i n platform web and e x c e s s i v e p o r o s i t y , T o a v o i d any c h a n c e o f a d d i t i o n a l platform f a i l u r e s , a l l re- maining blades were reworked (as shown i n F i g u r e 9B) t o r e d u c e t h e F i g u r e 9A. S k e t c h I l l u s t r a t i n g F i g u r e 9B. Sketch I l l u s t r a t i n g Flaw D i s c o u r a g e r Flow D i s c o u r a g e r F a i l u r e i n Blade N27 R e w o r k , During Second 50-Hour T e s t .

1 8 stress i n t h e f l o w d i s c o u r a g e r r e g i o n . The reworked b l a d e s were t h e n reassembled i n t o t h e d i s k w i t h b l a d e s s u b s t i t u t e d a c c o r d i n g to t h e s c h e d u l e . The rotor was r e b a l a n c e d , t h e damaged LP t u r b i n e components r e p l a c e d , and t h e e n g i n e reassembled for t h e t h i r d 50- test.

hour e n d u r a n c e The t h i r d 50-hour test c y c l e s i m u l a t e d a t y p i c a l m i s s i o n f o r a commuter a i r c r a f t ( F i g u r e 3 ) . N o problems were e n c o u n t e r e d d u r i n g t h i s test, and no s i g n i f i c a n t changes i n o p e r a t i n g parameters were o b s e r v e d as a r e s u l t o f t h e b l a d e r e w o r k . Post-test i n s p e c t i o n found no s i g n s o f distress i n t h e b l a d e s or any o t h e r component.

Blade s u b s t i t u t i o n s were a g a i n made a c c o r d i n g to t h e s c h e d u l e , t h e wheel r e b a l a n c e d , and t h e e n g i n e reassembled f o r t h e f o u r t h and f i n a l 50-hour test .

The f o u r t h 50-hour test c y c l e , d e s i g n e d t o p r o v i d e a n e v a l u a - t i o n of t h e l o w - c y c l e - f a t i g u e c a p a b i l i t y of t h e SC b l a d e s , is shown i n F i g u r e 4 . Normal a c c e l e r a t i o n s and d e c e l e r a t i o n s were main- t a i n e d f o r a l l c y c l e s . N o s i g n i f i c a n t e n g i n e test problems were e n c o u n t e r e d d u r i n g t h i s test, and e n g i n e performance remained con- post-test i n s p e c t i o n d i d n o t s i s t e n t w i t h p r e v i o u s t e s t i n g , The SC blades or any o t h e r hardware.

r e v e a l any d i s c r e p a n c i e s i n t h e F i g u r e s 10 and 11 show t y p i c a l l e a d i n g and t r a i l i n g edge v i e w s of b l a d e s tested f o r t h e e n t i r e 200 hours.

Detailed e x a m i n a t i o n s of sample test blades were accomplished a c c o r d i n g to t h e s c h e d u l e d e v a l u a t i o n . R e s u l t s of t h e s e i n s p e c - t i o n s are p r e s e n t e d i n t h e post-test e v a l u a t i o n s e c t i o n of t h i s report .

81674-4 Figure 1 0 . Leading-Edge V i e w of T y p i c a l Blades Subjected to t h e E n t i r e 200 Hours of Endurance T e s t i n g .

MP-87769 81674-3 F i g u r e 11. T r a i l i n g - E d g e V i e w of T y p i c a l Blades S u b j e c t e d to the E n t i r e 200 Hours of E n d u r a n c e T e s t i n g .

MP-87776 2 1 .1 I n t e r i m - T e s t Blade E v a l u a t i o n A f t e r 100 h o u r s of t e s t i n g ( C y c l e s A and B ) , r o u t i n e examina- t i o n o f t h e H P t u r b i n e rotor i n d i c a t e d t h a t one of t h e SC NASAIR 100 blades (S/N N27) had lost part of t h e a f t f l o w d i s c o u r a g e r . The weight o f t h e m i s s i n g piece was estimated a t a b o u t one gram. T o i d e n t i f y t h e c a u s e of t h i s f l o w d i s c o u r a g e r f a i l u r e and t o document t h e m i c r o s t r u c t u r e of t h e a i r f o i l , t h i s blade was v i s u a l l y and m e t a l l u r g i c a l l y examined. These e x a m i n a t i o n s ( F i g u r e s 1 2 and 1 3 ) r e v e a l e d t h a t t h e flow d i s c o u r a g e r web t h i c k n e s s was s i g n i f i c a n t l y below t h e minimum specified drawing t h i c k n e s s and o n l y a b o u t 65 p e r c e n t of t h e a v e r a g e web t h i c k n e s s . The p o r t i o n o f t h e p l a t f o r m web t h a t f a i l e d also had v e r y h i g h amounts o f o x i d i z e d ( s u r f a c e c o n n e c t e d ) porosity.

FPI and v i s u a l i n s p e c t i o n o f t h e remaining b l a d e s i n d i c a t e d t h a t comparable p l a t f o r m p o r o s i t y was p r e s e n t i n many o f t h e b l a d e s . However, no other blades had t h e e x c e s s i v e p l a t f o r m p o r o s i t y and below minimum web t h i c k n e s s t h a t combined to produce a f a i l u r e .

M i c r o s t r u c t u r a l a n a l y s i s o f blade S / N N27 a t o t h e r l o c a t i o n s on t h e blade w a s also performed ( F i g u r e s 1 3 and 1 4 ) . T h i s a n a l y s i s i n d i c a t e d t h a t t h e SC NASAIR 100 b l a d e had been s o l u t i o n h e a t - treated a t t h e upper end o f t h e s o l u t i o n h e a t - t r e a t m e n t window.

The e u t e c t i c gamma prime p h a s e was c o m p l e t e l y s o l u t i o n e d and minor i n c i p i e n t m e l t i n g was o b s e r v e d i n t h e blade root e S u b s t a n t i a l l y less m i c r o p o r o s i t y was n o t e d i n t h e a i r f o i l t h a n i n t h e blade root.

Energy d i s p e r s i v e X-ray ( E D X ) a n a l y s i s was used t o v e r i f y t h e NASAIR 100 c o m p o s i t i o n o f blade S/N N27.

PRECEDING PAGE BLANK NOT F

K

Visual Examination of SC NASAIR 100 Blade (S/N N 2 7 ) Figure 12.

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 75 percent of indicated magnifications.)

MP-87775 Figure 13, Extensive Interconnected and Oxidized Porosity (top) and Incipient Melting (bottom) Were Present in.the f NASAIR 10’0 Adjacent to the Failed Platform Web (SC NASAIR 100 Blade S/N N27), Photographs are 85 percent of indicated magnifications, MP-90751 A i r f o i l of Blade S/N N27 E x h i b i t e d a Figure 3.4, Microstructure T y p i c a l of F u l l y Processed S C NASAIR 100, (Photographs a r e 80 p e r c e n t of i n d i c a t e d m a g n i f i c a t i o n s . ) Mp-07770 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 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 indi- cates 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 metal- lurgical 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 V i s u a l Appearance of S C NASAIR 1 0 0 (S/N N 2 0 , N33, F i g u r e 15.

N48) and SC Alloy 3 (S/N A42) Blades a f t e r t h e 200-Hour Engine T e s t . V e r t i c a l Lines Indicate Location of L o n g i t u d i n a l Metallographic S e c t i o n .

MP-87838

e

t

Mp-90749 been completely solutioned, and some indications of incipient melt- ing were observed in the three blades examined. Results of the microstructural examination of Blade S/N N48 illustrate this condi- tion (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, how- ever, 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 t h t 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 loca- tion 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 air- foil 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 infor- mation, the Mu phase did not adversely affect the integrity of the SC NASAIR 100 KP turbine blades during the 200-hour engine test.

C A Microstructure of S C NASAIR 1 0 0 HP T u r b i n e B l a d e F i g u r e 1 7 .

(S/N N 4 8 ) after 200-Hour E n g i n e T e s t .

Mp-90750 MA16568 Figure 18, Sca i c r o s c o p e Photographs (2000X) of rbine Blade a f t e r 200-Hour Engine T e s t (S/N N48), Figure 19. Appearance of the B l o c k y Alpha-Tungsten Phase of SC NASAIR 100 HP (Arrows) in Airfoil Section Turbine Blade.

Mp-87765 F i g u r e 20. SEM Photograph and EDX A n a l y s i s of t h e Tungsten+ N i c k e l R i c h M u Secondary Phase i n t h e A i r f o i l S e c t i o n of an SC NASAIR 100 B l a d e (S/N N 4 8 ) a f t e r 200 Hours of S u c c e s s f u l Engine T e s t i n g .

MP-87772

SECTION V 5.0 CONCLUSIONS The f o l l o w i n g c o n c l u s i o n s were based upon t h e results o f t h e e n g i n e t e s t i n g and p o s t - t e s t a n a l y s i s of t h e SC t u r b i n e b l a d e s pre- s e n t e d i n t h i s report: 0 I n c o r p o r a t i n g t h e MATE H P t u r b i n e b l a d e s and s u p p o r t hardware i n t o t h e test e n g i n e reduced TSFC by 1 . 4 p e r c e n t and T5 by 2.5 p e r c e n t compared to a b a s e l i n e e n g i n e w i t h d i r e c t i o n a l l y s o l i d i f i e d H P t u r b i n e b l a d e s .

0 The MATE s i n g l e - c r y s t a l t u r b i n e b l a d e s , b o t h NASAIR 100 and A l l o y 3, showed no s i g n s of d i s t r e s s a f t e r success- f u l l y c o m p l e t i n g o v e r 200 h o u r s of endurance e n g i n e t e s t i n g .

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 invest-

ment 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 equip- ment, thus eliminating one advantage of the exothermic process. Therefore, unless the proprietary aspect of the SC withdrawal process becomes a significant problem, fur- ther development of the exothermic process i s not recom- mended.

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 REFERENCES 1 . Strangman, T.E., B.R. Heath, and M . Fujii: Low-Cost, Single- Crystal 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 Single-

Crystal Superalloy," in Superalloys 1980 - Proceedings of the

Fourth International Symposium on Superalloys, Champion, Pennsylvania, September 1980, p p : 5 1 3 - 5 2 0 .

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.

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Attn C. Hoelzer C32-5 Attn W . Freeman Bethpage, NY 11719 475 Steamboat Road Greenwich, CT 06830 Cessna Airctaft Co. Chromoloy Attn C. Gonzales Attn D. Kenton (Turboprop) P , O . Box 10699 Wallace Division Midwest City, OK 73140 P.O. Box 7704 Wichitar KS 67277 Colt Industries Colt Industries Attn E. Dulis Attn 3 . Moll Crucible-Inc, Mat'ls Rsch Crucible Inc, Mat'ls Rsch P. 0. Box 88 P: 0. Box 88 Pittsburgh, PA 15230 Pittsburgh, PA 15230 Delta Airlines Douglas Aircraft Co.

Attn W.J. Overend Attn Library Atlanta Airport McDonnell Douglas Corp Atlanta, 43A 30320 3855 Lakewood Blvd.

Long Beach, CA 90846 Ford Mota? Company Ford Motor Company Attn C . Feltner Attn Y. Telang Metallurgical Dept. Metallurgical Dept.

P.O. Bo% 2053 P.O. Box 2053 Dearborn, Mf 98121 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.

D ~ S T ~ ~ B U T I Q N LIST (CQNTD) Gould In@. Gould Inc.

, Attn E . N. A ua Attn G . Davies

5 4 0 E. 105th St.

5 4 0 E. 105 S f Cleveland, OH-49106 Cleveland, OH 44108 Howmet Corporation Howmet Corporation Technical Center Technical Center Attn J. VanderSluis ,Attn L. Dardi 699 Benston Rd. 699 Benston Rd.

Whitehall, MI 49461 Whitehall, MI 49961 Howmet Corporation Howmet Corporatidn Turbine Components Div. Turbine Components Div.

Attn Mr. Burd Attn E. Carozza Dover, NJ 07801 Mr. Burd Dover, NJ 07801 Huntington Alloys Div. IIT Research Institute International Nickel Co. Attn M. A. Howes Attn J. Cundiff 10 West 35th Street Huntington, W V 25720 Chicago, IL 60616 Intermetco jetshapes, Inc.

Attn J. Siergiej Attn F. Jaeger 300 Concord Road Rockleigh Industrial Perk Wayland, MA 01778 Rockleigh, NJ 07647 Kelsey Hayes Co. Thomas McNamara Attn Wm. E. Good Consultant Heintz Div. 42701 Main St.* SP 154 Front St. and Olney h e . Sen Jacinto Philadelphia, PA 19120 CA 92383 METCO Inc. METCO Inc.

Attn V. Lanza Attn C. Lewis 1101 Prospect Ave. 3 4 0 0 A Oak Cliff Rd.

Westbury, L.I., NY 11590 Atlanta, GA 30340 Rand Corp. Rocketdyne Division Attn J. Richard Nelson Attn J. Frandsen Washington Research Div. Rockwell International 2100 M St. 6633 Canoga Ave.

Washington, DC 20037 Canoga Park, CA 91364 5lSTRlBUTlON LIST (CONTD) Rockwell International Special Metals, Stla, Attn L.L. McHughes Attn J. Pridgeon General Aviation Div. Middle Settlement Road 5001 N. Rockwell Ave.

New Hartford Bethany, OK 73008 Neu York 13413 Huntington A110 s Div. Huntington Allo s Div.

International Kickel Co.

International Kickel Co.

Attn R. Haeberle, 3r Attn F. Perry Huntington, W V 25720 Huntington, W V 25720 IIT Research Institute IIT Research Institute Attn V. Hill Attn. R. Firestone 1 0 West 39th Street 10 West 35th St.

Chicago, IL 60616 Chicago IL 60616 Kaman Sciehces Corp.

Kelsey Hayes Go.

Attn D . H. Seita Attn W.G, Koby P. 0. Bow 7463 Heinta Div.

Colorado Springs Front St. and Olney Ave.

60 80933 Philadelphia, PA 19120 MCIC Martin Marietta Battelle Memorial Inst. Attn C.H. Lund Attn H. Mihdlin 15 N. Windsor Rd.

COlUmbUSr Ofl 93201 Arlington Hts IL 6 0 0 0 4 Pacifia Southwest Airlines Pan American Woi.16 lirways Attn E . Chambers Attn W.B. Hibbs 7707 Consolidated Way Pan Am Buildin San Diego, CR 92121 New York. NY f o o l 7 Rockwell International Rockwell International Attn V. Ta lor Attn. E . Mathwi General Avration Div. ' General Rviation #iv .

5 0 0 1 N. Rockwell Ave.

5001 N. Rockwell Ave.

Bethany, OK 73008 Bethany, OK 73008 Cabot Cor Stellite Division Attn D . tiaistrom Cabot Corporation 1020 W , Park Xve. Attn M. Rothman Kokomar I N 46901 1020 West Park he.

Kokomo, IN 46901 ~ ~ S T R I B U T ~ O N LIST (CONTD) Sundstrand Swearingen Aviation Cotp.

Attn D. Augustine Attn R. McKelvey 4747 Harrison Ave P.O. Box 32486 Rockford, IL 61101 San Antonio, TX 78284 TRW Inc . TRW Inc .

Attn C.S. Kortovich Attn J . McCarthy 23555 Euclid Ave.

23555 Euclid Ave.

Cleveland, OH 44117 Cleveland, OH 4411) TRW Inc . .

TRW Inc Attn T. Piwonka Attn C. Vishnevskv 23555 Euclid Ave. 23555 Euclid Ave.- Cleveland, OH 44117 Cleveland, OH 44ll# Turbine Support Co.

Union Carbide Corp.

Attn R. Van Cleaf eLinde Division P. 0. Box 20148 Attn: E. B. Cook, Jt.

P.O. Box 20148 441 Sackett Pt. Rd.

San Antonio, TX 78220 North Haven, CT 06475 United Tech Rsch Center United Tech Rsch Center Attn B. Thomson Attn Library East Hartford, CT 06108 E a s t Hartford, C t bSlOb Universal Cyclops Universal Cyclops Attn R. Gasior Attn Wm. Kent Mayer Street Player St Bridgville, PA 15017 Bridgeville PA 15017 Vought Corp. Westinghouse R + t, Centei: Attn D. Moon Attn Library 2-50370/TL 7-67 Beulah Rd.

Pittsburgh, PA 15235 P.O. Box 5907 Dallas TX 75222 Teledyne Allvac Teledyne Allvaa Attn R. Kennedy Attn S. C. Pearman P. 0. Box 759 Wm. Thomas P. 0. Box 759 Monroe, NC 28110 Monroe, NC 28110 DIST~IBUTIQNLIST (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 Player St.

East Hartford, CT 06108 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 Bolt 10864 P,O, Box l086Q P,O.

Pxttsbutgh, PA 19236 Pittsburgh, PA 15236 Westinghouse Electric Co. , Attn E. Crombie C-210 P.O. Box 251 Concordville, PA 19331 5 4 1. Repart No 2. Government Accession No. 3. Recipient's Catalog No.

CR-17 4 6 52 4. Title and Subtitle April 1984 Low-Cost S ing le-Cry stal Turbine Blades 6. Performing Organization Code llolume 2 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 11. Contract or Grant No A Division of The Garrett Corporation Phoenix, Arizona 85010 NAS3-20073 13. Type of Repon and Period Covered Project Completion Report 12. Sponsoring Agency Name and Address Project 3 National Aeronautics and Space Administration 14. Sponsoring Agency Code Washington, D.C. 20546 15. Supplementary Notes Robert L. Dreshfield, Materials Division Project Manager: NASA-Lewis Research Center, Cleveland, Ohio 16. Abstract The overall objectives o f Project 3 were to develop the exothermic casting pro- cess 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 endur- ance engine test ing.

These Program objectives were achieved. The exothermic casting process was successfully developed into a low-cost nonproprietary method for producing single- crystal 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 18. Distribution Statement Columnar -Gr a in Tur b ine-Blade S ing le-Cr y s t a1 MAR-M 247 Alloy-Selection NASAIR 100 Ni-base Alloys Alloy 3 Investment Casting Exothermic-Heatinq 19. Security Classif. (of this report) I 20. Security Classif. (of this page) 1 21. NO. of Pages I 22. Rice' Unclassified Unclassified 61 NASA-(2-168 (Rev. 10-75)