Energy efficient engine high-pressure turbine supersonic cascade technology report

TABLE O F CONTENTS Page Section 1.0 Sumnary 2.0 I n t r o d u c t i o n 3.0 Vane Cascade Program 3 . 1 Analysis and Design 3.2 Fabrication and Assemb l y 3.3 Testing 3.3.1 General Description 3.3.2 Test Faci l i t y and Instrumentation 3.3.3 Test Procedures 3.3.4 Performance Test P i an 3.3.5 Data Reduction and Analysis 3.3.6 Experimental Uncertainty 3.4 Results 3.4.1 F low Visualizations 3.4.2 P e r i o d i c i t y Evaluation 3.4.3 Pressure D i s t r i b u t i o n Analysis 3.4.4 Umooled Cascade Results 3.4.5 Coo led Cascade Resu I t s f o r S t r a i g h t M a l l Cascade 3.4.6 S m a r y o f Vane Cascade Resu i t s 4.0 B lade Cascade Progran 4.1 Analysis and Design 4.2 Fabrication and Assembly 4.3 Testing 4.3.1 General Description 4.3.2 Test Faci l i t y and Instrumentation 4.3.3 Test Procedures 4.3.4 Performance Test Plan 4.3.5 Data Reduction and Analysis 4.3.6 Experimental Uncertainty 4 . 4 Results 4 . 4 . 1 F low Visualizations 4.4.2 P e r i o d i c i t y Evaluation 4.4.3 Uncoo led Cascade Resu I t s 4.4.4 Coo led Cascade Resu I t s 4.4.5 Sumnary of Blade Cascade Resu I t s 5.0 Conc lusions 5.1 Vane Cascades 5.2 B lade Cascades

TABLE O F CONTENTS (Cont'd) Appendix A: Tab l e A - 1 Vane A i r f o i l Coordinates f o r S-Wal l Cascade Table A-2 Vane A i r f o i l Coordinates f o r S t r a i g h t Wall Cascade Tab l e A-3 Base B lade A i r f o i 1 Coordinates Table A-4 Overcambered Blade A i r f o i l Coordinates Tab l e A-5 Straightback B lade A i r f o i l Coordinates Appendix B: L i s t o f Symbols References Figure 1 Supersonic Cascade Program Logic D i agram 2 Schematic o f Turbine I n l e t Vane Endwal l Contour S t r a i g h t Wall and S-Wall Cascade Configurations 4 Vane Cascade Geometry Comparison of Pressure D i s t r i b u t i o n s 6 Cascade Configuration H i gh-Pressure Turbine B lade Cooling Model High-Pressure Turbine Assembled S t r a i g h t Wall Vane Cascade Before I n s t a l l a t i o n i n Test Tunne l High-Pressure Turbine Vane Cascade Test Faci l i t y High-Pressure Turbine Vane Cascade Endwal l S t a t i c Pressure Tap Locations High-Pressure Turbine S t r a i g h t Wall Vane Cascade S t a t i c Pressure Tap Locations High-Pressure Turbine S-Wa l l Vane Cascade S t a t i c Pressure Tap Locations S t r a i g h t W a l l Cascade F low Visua li z a t i o n Viewed from Upstrean S-Wal l Cascade F low V i s u a l i z a t i o n Viewed from Downstream Cascade Secondary F low Schematic Cascade F low Passages S-Wa l l Cascade Mid-Span Pressure D i s t r i b u t i o n P e r i o d i c i t y Evaluation S t r a i g h t Wal l Mid-Span Pressure D i s t r i b u t i o n P e r i o d i c i t y Evaluation

S t r a i g h t - W a l l Cascade - Comparison o f Cross-Channe l Data

With Two-Dimensiona l Potent i a l F low Computation L I ST O F ILI-USTRATIONS (Cont d. ) Figure Page

S-Wa l l Cascade - Comparison o f Cross-Channe l Data w i t h

Three-Dimensional I n v i s c i d Flow F i e l d Computation a t S t r a i g h t Wal l End o f Cascade

S-Wa l l Cascade - Comparison o f Spanwise Data a t P lanes A

and B w i t h Three-Dimensional I n v i s c i d Flow F i e l d Computation

S- W a i l Cascade - Comparison o f Spanwise Data a t P lanes

C and D w i t h Three-Dinensional I n v i s c i d Flow F i e l d Computation

S-Wal l Cascade - Comparison o f Cross-Channel Data w i t h

Three-Dimensiona l I n v i s c i d F low F i e I d Computation a t S-Wa l l End o f Cascade S t r a i g h t Wal l Cascade Pressure Loss Contours S-Wall Cascade Pressure Loss Contours Spanwise D i s t r i b u t i o n o f T o t a l Pressure Loss f o r S-Wal l Cascade Spanwise D i s t r i b u t i o n o f T o t a l Pressure Loss f o r S t r a i g h t W a l l Cascade Spanwise D i s t r i b u t i o n of E x i t Yaw Angle f o r t h e S t r a i g h t W a l l Cascade Spanwise D i s t r i b u t i o n o f E x i t Yaw Angle f o r the S-Wall Cascade Comparison of S t r a i g h t W a l l Cascade Data w i t h t h e Corre l a t i o n o f M o r r i s and Hoare Comparison of S-Wal l Cascade Data w i t h the ' P r o f i l e A t C o r r e l a t i o n o f M o r r i s and Hoare S t r a i g h t Wall Cascade Midspan Pressure Loss vs Mach Number S t r a i g h t W a l l Cascade Predicted and Measured Pressure

D i s t r i b u t i o n s - Mach Number = 0.837

S t r a i g h t W a l l Cascade Predicted and Measured Pressure

D i s t r i b u t i o n s - Mach Number = 0.877

S t r a i g h t W a l l Cascade Predicted and Measured Pressure

D i s t r i b u t i o n s - Mach Number = 0.906

CI ST OF ILLUSTRATIONS (Cont ' d. )

F i g u r e P a g e S t r a i g h t W a l l Cascade P r e d i c t e d and Measured Pressure 3 6

D i s t r i b u t i o n s - Mach Number = 0.974 40

Straight-Wa 1 l Cascade Midspan E x i t A i r Ang l e vs Mach N urnb e r 40 Pressure D i s t r i b u t i o n vs Chord a t Three Span L o c a t i o n s Spanwise D i s t r i b u t i o n o f T o t a l Pressure Loss f o r t h e S t r a i g h t W a l l Cascade Spanwise D i s t r i b u t i o n o f E x i t Yaw Angle f o r t h e S t r a i g h t Wal l Cascade 4 3 Pressure Loss E f f e c t s Due t o T r a i l i n g Edge I n j e c t i o n a t 4 5 Design P o i n t Coolant Flow E x i t A i r Angle E f f e c t s Due t o T r a i l i n g Edge I n j e c t i o n a t Design P o i n t Coolant F l o w 45 4 6 Vane Cascade T r a i l i n g Edge F low C a l i b r a t i o n Pressure Loss E f f e c t s Due t o Pressure Surface Coolant I n j e c t i o n 48 E x i t A i r Ang le E f f e c t s Due t o Pressure Surface Coolant I n j e c t i o n 4 8 Pressure Loss E f f e c t s Due t o Pressure Surface Coolant 4 9 F low I n j e c t i o n Rate V a r i a t i o n E x i t A i r Ang le E f f e c t s Due t o Pressure Surface Coolant 4 0 Flow I n j e c t i o n Rate V a r i a t i o n Pressure Loss E f f e c t s Due t o Suction Surface Coolant I n j e c t i o n E x i t A i r Angle E f f e c t s Due t o Suction Surface Coolant I n j e c t i o n 50 Vane Suction Surface I n j e c t i o n Midspan Loss vs Coo l i n g F low Rate 5 1 S t r a i g h t Wal l Cascade E x i t A i r Angle vs C o o l i n g Flow Rate

L I V O F ILLUSTRATIONS (Cont Id. )

F i g u r e

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52 S t r a i g h t Wall Cascade Pressure Loss vs S u c t i o n Surface Coolant Flow Rate a t E x i t Mach Number Equa! t o 0.9 E f f e c t s of Changes i n S u c t i o n Surface Coolant F low Rate and E x i t Mach Number cn Cascade Pressure Loss S t r a i g h t W a l l Cascade M i dspan Loss vs Mach Number With and Without Design P o i n t Coolant Flow I n j e c t i o n Measured E x i t A i r Ang l e as a F u n c t i o n o f S u c t i o n Surface Coolant Flow V a r i a t i o n f o r E x i t Mach Numbers of 0.83 and 0.9 High-Pressure T u r b i n e Base Blade Design Parameters and P r e d i c t e d Pressure D i s t r i b u t i o n High-Pressure Turbine Overcambered Blade Design Parameters and P r e d i c t e d Pressure D i s t r i b u t i o n High-Pressure T u r b i n e S t r a i g h t b a c k Blade Design Parameters and P r e d i c t e d Pressure D i s t r i b u t i o n High-Pressure Turbine Base Blade T r t ~ i l i n g Edge Coolant E j e c t i o n Geometry Photograph cf Instrumented B I ade Schematic of U n i t e d Techno logies Research Center Variab l e D e n s i t y Supersonic Cascade Wind Tunne l Schematic o f Cascade I n s t a l l a t i o n Tunne l Test Section w i t h a Cascade I n s t a l led Schematic c f S c h l i e r e n System High-Pressure Turbine Blade Cascade S t a t i c Pressure Tap L o c a t i o n s I s o m e t r i c View o f Wedge Probe T i p R e l a t i o n s h i p Between Actual E x i t Mach Number and I s e n t r o p i c E x i t Mach Number f o r t h e Base A i r f o i l Cascade R e l a t i o n s h i p Between Actual E x i t Mach NumDer and I s e n t r o p i c E x i t Mach Number r o r t h e Overcamber A i r f o i 1 Cascade L I S T O F ILLUSTRATIONS (Cont ' d. ) F i g u r e

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69 Re l a t i o n s h i p Between Actua l E x i t Mach Number and I s e n t r o p i c E x i t Mach Number f o r t h e S t r a i g h t b a c k A i r f o i I Cascade 7 1 Re l a t i onship Between Actua l E x i t Mach Number and I s e n t r o p i c E x i t Mach Number f o r t h e Base A i r f o i l Cascade w i t h T r a i l i n g Edge Coo l a n t F low Discharge 7 1 F low Visua l i z a t i o n Photograph o f Overcambered B lade S u c t i o n Surface f o r an I s e n t r o p i c E x i t Mach Number o f 1.3 74

Sch l i e r e n Observation of Ob l i q u e Shock Wave - Boundary Layer

I n t e r a c t i o n on B lade Suction S u r f ace 75 D e p i c t i o n of Shock Boundary Layer I n t e r a c t i o n 7 6 Loss f o r I n d i v i d u a l B6se A i r f o i I s I n Supersonic and Subsonic F low Sch l i eren Photograph Showing F low S t r u c t u r e w i t h 2 e f l e c t e d Shocks 77 Spanwise V a r i a t i o n o f T o t a l Pressure Loss f o r Base A i r f o i l a t 0.65, 0.79 and 1.3 E x i t Mach Numbers 7 8 Uncoo led Cascade Test (Base B l ade Sect i o n ) Pressure D i s t r i b u t i o n a t 0.599 E x i t Mach Number Urlcoo led Cascade T e s t (Base B lade Section) Pressure D i s t r i b u t i o l ~ a t 0.763 E x i t Mach Number Uncoo led Cascade Test (Base Blade Section) Pressure D i s t r i b u t i o n a t 0.968 E x i t Mach Number Jncoo led Cascade Test (Base B l ade Sect i o n ) Pressure D i s t r i b u t i o n a t 1.044 E x i t Mach Number Uncoo led Cascade Test (Base B lade Sc i o n ) Pressure D i s t r i b u t i o n a t 1.137 E x i t Mach NUIT,, Uncoa led Cascade Test (Base Blade Section) Pressure D i s t r i b u t i o n a t 1.229 E x i t Mach Number Uncoo led Cascade Test (Base B lade Section) Pressure D i s t r i b u t i o n a t 1.325 E x i t Mach Number Uncoo led Cascade Test (Base B l ade Section) Pressure

D i s t r i b u t i o n a t 1.251 E x i t Mach Number - I n l e t Gas

Angle o f 33 degrees

L I ST O F ILLUSTRATIONS (Cont ' d. )

Page F i g u r e 85 Uncoo led Cascade Test (Base B lade Section) Pressure

D i s t r i b u t i o n a t 1.239 E x i t Mach Number - I n l e t Gas

Ang le o f 58 degrees Sch I i e r e n Photograph of Base B lade Mean Sect i o n a t 1.1 I s e n t r o p i c E x i t Mach Number Sch l i e r e n Photograph o f Base Blade Mean Section a t 1.2 I s e n t r o p i c E x i t Mach Number Schlieren Photograph o f Base Blade Mean Section a t 1.27 I s e n t r o p i c E x i t Mach N~mber 8 5 Sch l i e r e n Photograph o f Base B iade Mean Section a t 1.4 I s e n t r o p i c E x i t Mach Number I n d i v i d u a l A i r f o i l T o t a l Pressure Loss vs E x i t I s e n t r o p i c Mach Number Base Blade I n d i v i d u a l A i r f o i l Base Pressure vs E x i t I s e n t r o p i c Mach Number Base B 1 ade 87 I n d i v i d u a l A i r f o i l E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Number Base B l ade 87 Uncoo led Cascade Test Overc~mbered B lade Sect i o n Pressure D i s t r i b u t i o n a t 0.624 E x i t Mach Number 89 Uncoo led Cascade Test Overcambered B lade Section Pressure D i s t r i b u t i o n a t 0.761 E x i t Mach Number 89 Uncooled Cascade Test Overcambered Bladk Section Pressure D i s t r i b u t i o n a t 0,958 E x i t Mach Number 89 Uncoo led Cascade Test Over cambered B lade Sect i o n Pressure D i s t r i b u t i o n a t 1.008 E x i t Mach Number 89 Uncoo led Cascade Test Overcambered B lade Sect i m Pressure D i s t r i b u t i o n a t 1.122 E x i t Mach Number 90 Uncooled Cascade Test Overcambered Blade Section Pressure D i s t r i b u t i o n a t 1.226 E x i t Mach Number 90 Uncoo led Cascade Test Overcambered B lade Sect i o n Pressure D i s t r i b u t i o n a t 1.335 E x i t Mach Number 40 L I ST OF ILLUSTRATIONS (Cant ' cl. ) Page F i gure 100 Sch li eren Photograph o f Overcambered B 19de Mean Sect i o n a t 1.08 I s e n t r o p i c E x i t Mach Number 9 1 Sch l i eren Pnotograph o f Overcambered B lade Mean Sect i o n a t 1.19 I s e n t r o p i c E x i t Mach Number 92 Sch li eren Photograph o f Overcambered B lade Mean S e c t i o n a t 1.28 I s e n t r o p i c E x i t Mach Number 9 3 Sch li eren Photograph o f Overcambered B lade Mean Section a t I-. 39 I s e n t r o p i c E x i t Mach Number 94 I n d i v i d u a l A i r f o i l Mass Averaged Tota l Pressure Loss vs E x i t I s e n t r o p i c Mach dumber 95 I n d i v i d u a l A i r f o i l Base Pressure vs E x i t I s e n t r o p i c Mach Number 95 I n d i v i d u a l A i r f o i l E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Number Overcambered B l ade 9 5 Uncoo led Cascade Test Straightback B lade Section Pressure D i s t r i b u t i o n a t 0.624 E x i t Mach Number 97 Uncoo led Cascade Test Straightback B lade Sect i c n Pressure D i s t r i b u t i o n a t 0.782 E x i t Mach Number Uncooled Cascade Test Straightback Blade Section Pressure D i s t r i b u t i o n a t 0.956 E x i t Mach Number 97 Section Pressure Uncoo led Cascade Test Straightback B lade D i s t r i b u t i o n a t 1.042 E x i t Mach Number Uncoo led Cascade Test Straightback B lade Section Pressure D i s t r i b u t i o n a t 1.151 E x i t Mach Number 98 Uncoo led Cascade Test Straightback B lade Sect i o n Pressure D i s t r i b u t i o n a t 1.237 E x i t Mach Number 98 Uncoo led Cascade Test Straightback B lade Sect i o n Pressure D i s t r i b u t i o n a t 1.S42 E x i t Mach Number 98 Sch I i e r e n Photograph o f Straightback B lade Mean Section a t 1.10 I s e n t r o p i c E x i t Mach Number 99 L I ST OF ILLUSTRATIONS (Cont d. ) F i gure Page 115 Sch l i eren Photograph o f Straightback B l ade Mean S e c t i o n a t 1.19 I s e n t r o p i c E x i t Mach Number 100 Sch l i eren Photograph o f Straightbac'c S lade Mean S e c t i o n a t 1.29 I s e n t r o p i c E x i t Mach Number 101 Sch l i eren Photograph o f Straightback B l ade Mean S e c t i o n a t 1.35 I s e n t r o p f c E x i t Mach Number 102 I n d i v i dua I A i r f o i l Mass Averaged Tota l Pressure Loss vs E x i t I s e n t r o p i c Mach Number 103 I n d i v i d u a l A i r f o i l Base Pressure vs E x i t 1sent:opic Nach Number 103 I n d i v i d u a l A i r f o i l Mass Averaged E x i t A i r Angle vs E x i t I s e ] t r o p i c Mach Number Cooled Cascade Test Base Glade Section Pressure D i s t r i b u t i o n a t 0.565 E x i t Mzch Number 105 Coo led Cascade Test Base B lade Section Pressure D i s t r ; $ u t i o n a t 0.62 F v i t Mach Number 105 Cooled Cascade Test Base Blade S e c t i o n Presrure D i s t r i b u t i o n a t 0,765 E x i t Mach Number Cooled Cascade Test Base Blade Section Pressure !C5 D i s t r i b u t i o n a t 0.772 E x i t Mach Lumber Coo led Cascade Test Base B lade S e c t i o n Pressure D i s t r i b u t i o n a t 0.98 E x i t Mach Number 106 Coo led Cascade Test Base B lade Section Pressure Dist:ibution a t 0.989 E x i t Mach Number Coo led Cascade Test Base B lade S e c t i o n Pressure D i s t r i b u t i o n a t 1,026 E x i t Mach Number Coo led Cascade Test Base B lade S e c t i o n Pressure D i s t r i b u t i o n a t 1.034 E x i t Mach Number 136 1-1 ST O F ILI-USTRATIONS (Cont ' d. ) Page -- Coo led Cascade Test Base B lade Sect i o n Pressure D i s t r i b u t i o n a t 1.034 E x i t Mach Number Coo led Cascade Test Base B lade Section Pressure D i s t r i b u t i o , i a t 1.14 E x i t Mach Numher Cooled Cascade Test Base B lade Section Pressure D i s t r i b u t i o n a t 1.157 :,,it Mach Number Coo led Cascade T e s t 8ase B lade Sect i o n Pressure D i s t r i b u t i o n ? : 1.263 E x i t Mach Number Coo led Cascade Test Base B lade Section Pressure D i s t r i b u t i m a t 1.273 E x j t Mach Nufiber Cooled Cascade Test Base Blade Section Pressure D i s t r i b u t i o n a t 1.274 E x i t Mach Number Cooled Cascade Test Base Blade Section Pressure D i s t r i b u t j o n a t 1.296 E x i t Mach Number Sch l i e r e n Photograph o f Base B lzde Mean Section a t 1.3 I s e n t r o p i c E x i t Mach Number - Wc/Wm = 0 109 Sch l i e r e n Photograph o f Base Blade Mean Section a t 1.3 I s e n t r o p i c 1 1 0 E x i t Kach Nurnber - 'rlclWm = 0.012 Sch l i e r e n Photograph o f Base B lade Mean Section a t 1.3 I s e n t r o o i c 1.11

E x i t Mach Number - Wc/Wm = 0.024

Sch l i e r e n Photograph o f Base B la+? Mean Section a t 1.3 l s e n t r o p i c

E x i t Mach Number - Wc/Wm = 0.0272 112

Base Pressure C o e f f i c i e n t vs Plenum E x i t I s e n t r o p i c Mach Num5er a t Various Coo l i n g F lcw Rates 113 Base B lade T o t a l Pressure Loss VC, COO t i n 5 F low Ratio; MZi = 0.65 1.14 Base B lade T o t a l Pressure Loss vs Coo l i n g F l 9 w Ratio; M z i = 0.79 114 Base B lade Tota i Pressure Loss vs Coo l i n g F low R a t i o ; M 2 i = 1.00 114 LIST OF ILlUST9ATIONS (Cont ' d. ) F i gure Page 144 Base Blade Tota l Pressure Loss vs Coo l i n g F low Ratio; I t 5 M2i = 1.1 Base B lade T o t a l Pressure Loss vs Coo l i n g F low Ratio; MZi = 1.2 Base B lade Tota l Pressure l o s s vs Coo l i n g F low Rat ie; MZi 1 . 3 Schematic of Shock and Wake Loss Contributions t o T o t a l Pressure Loss Gapwise D i s t r i b u t i o n s o f T o t a i Pressure t o s s For Base A i r f o i I s With Severa I Coo li ng F low Rates And I s e n t r o p i c Mach Numbers o f 1.0 And 1.3 T o t a l Pressure, Shock Fnd Wake Loss vs E x i t I s e n t r o p i c Mach Number a t Various Coo l i n g F low Rates Schematic o f Shock P a ~ ~ e r n Formed a t Ribs Between Tra i l i ng Edge Coolant E j e c t i o v H J les Bas? Blade E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Num5er a t Various Coo i i n g F low Rates Conparison o f Total Pressure Loss vs E x i t I s e n t r o p i c Mach N u~nb er Comparison o f Base Pressure vs E x i t I s e n t r o p i c Mach Number Comparison o f E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Number Comparison o f Total Pressure Loss vs E x i t I s e n t r o p i c Mach Number Comparison o f E x i t A i r Angle vs E x i t I s e r ~ t r o p i c Mach Number Nomenclature Used t o Define A i r f o i l Coordinate, Nomenclature Used t o Define P r o f i l e d Wall Coordinates i n S- W a I I Vane Cascade The s i n g l e stage high-pressure t u r b i n e f o r t h e Ent,,y E f f i c i e n t Engine employs a i r - c o o l e d blades and vanes and advanced aerodynamic concepts such as contoured vane endwalls t o achieve i t s goal e f f i c i e n c y and l i f e . The t u r b i n e was designed t o have a high r a t i o o f wheel speed t o s p e c ' f i c work ( v e l o c i t y r a t i o ) a low r a t i o o f through-flow t o wheel t a n g e n t i a l v e l o c i t y (Cx/U) and h i g h AN^ (product o f annulus area and wheel speed squared). I n addition, design s t u d i e s i n d i c a t e d t h a t t u r b i n e e f f i c i e n c y could be f u r t h e r improved b y increasing t h e t b . b i n e r e a c t i o n l e v e l from a palanced Mach number design t o a design w i t h a subsonic vane and a supersonic blade, w i t h c a r e f u l a t t e n t i o n p a i d t o blade a i r f o i l curvature a f t o f the t h r o a t . Cooling i s achieved b y a conbination o f i n t e r n a l conductive and exte: n a l f i l m c o o l i n g techniques.

Testing o f s t r a i g h t endwall and contoirred etldwal 1 (S-wall) vane cascades i n - dicated t h a t the S-wall cascade had i7 percent l e s s full-passage, mass-averaged pressure loss than t h e s t r a i g h t w a l l cascade. I n t h e mid-span two-dimensional f l o w region o f the cascade. vane e x i t a i r angle was g e n e r a l l y i n s e n s i t i v e t o changes i n cndwall c o n f i g u r a t i o n as we1 1 as v a r i a t i o n s i n e x i t Mach number and coolant f l o w r a t e s . However, mid-span t o t a l pressure l o s s almost doubled when design-point coolant f l o w was e j e c t e d from a l l discharge ports. The l a r g e s t c o n t r i b u t o r was suction surface coolant f l o w i n j e c t i o n , which had a p e n a l t y about 5 times higher than t h a t due t o e i t h e r pressure surface o r t r a i l i n g edge coolant flow. V a r i a t i o n s i n e x i t Mach nunber confirmed t h a t t h e component vane design was f r e e o f transonic drag r i s e i n t h e range o f intended operating conditions.

Testing of blade cascades representing the high-pressure t u r b i n e component a i r f o i 1 base1 i n e design and two a l t e r n a t e d i s t r i b u t i o n s o f a i r f o i 1 c u r v a t u r e v e r i f i e d the a c c e p t a b i l i t y o f the component design. I n addition, base a i r f o i l pressure loss was r e l a t i v e l y i n s e n s i t i v e t o v a r i a t i o n s i n i n l e t ?as angle over a range o f 25 degrees, i n d i c a t i n g t h a t the component blade design has good incidence range. The a d d i t i o n o f t r a i l i n g edge coolant f l o w e j e c t i o n caused a s l i g h t increase i n base blade pressure loss a t subsonic e x i t Mach numbers, b u t reduced t o t a l pressure l o s s a t supersonic e x i t Mach nunbers r e l a t i v e t o t h e base blade without coolant f l )w e j e c t i o n . This was a t t r i b u t a b l e t o t h e f a c t t h a t flow e j e c t i o n reduced thtn t r a i l i n g edge shock s t r e n g t h such t h a t decreases i n shock losses predominated over corresponding increases i n wake mixing losses. Measured e x i t a i r angles were w i t h i n +1.0 t o -2.0 degrees o f t h e de- s i g n e x i t a i r angle o f 17 degrees a t the design p o i n t e x i t Mach nunber. T r a i l - i n g edge flow e j e c t i o n had l i t t l e e f f e c t on e x i t a i r angle.

Predicted a i r f o i l pressure d i s t r i b u t i o n s were i n good agreement w i t h measured d i s t r i b u t i o n s f o r b o t h t h e cooled and uncooled vane and blade cascades a t sub- sonic e x i t Mach nuribers. A t supersonic e x i t Mach nunbers, agreement f o r t h e blade cascades was f a i r due t o data s c a t t e r i n t h e a i r f o i l f l o w recompression region. This was a t t r i b u t a b l e t o non-periodic shocks i n the blade passage.

2.0 INTRODUCTION The objective of the NASA Energy Efficient Engine Development and Integration program i s to develop, evaluate, and demonstrate the technology for achieving lower instal led fuel consumption and lower operating costs i n future commercial turbofan engines. N A S A has set minimum goals of 12 percent reduction i n thrust

specific fuel consumption (TSFC) , 5 percent reduct ion in direct operating cost

(DOC), and 50 percent reduction i n performance degradation for the Energy Efficient Engine ( f l i g h t engine) relative to the JT9D-7A reference engine. In addition, environmental goals on emissions (meet the proposed Environmental Protection Agency 1981 regulation) and noise (meet Federal Aviation Regulations 36-1978 standards) have been established.

The Pratt & Whitney A i r ~ r . ~ t Energy Efficient Engine high-pressure turbine i s a single-stage design. A single-stage design has certain advantages when com- pared to i t s multi-stage counterpart. Single stage turbines require no inter- stage seal s, require fewer cooled airfoils, and contain fewer leakage paths.

The inherent design simplicity of the single stage reduces engine i n t i t i a l cost, maintenance material cost, and overall engine weight.

The purpose of the Energy Efficient Engine High-Presslire Turbine Supersonic Cascade Test Program was to (1) verify the benefits of vane endwall contouring in the high-pressure turbine component, ( 2 ) determine the performance penalties or benefits associated with the injection of coolant flow into the flow f i e l d surrounding the vane and blade airfoils, and (3) verify that the distribution of curvature selected for the component blade airfoil geometry achieves design performance objectives. An additional objective was to employ measured data to assess the accuracy of analytical methods and to gain a better understanding of the flow f i e l d within the cascade, particularly as i t i s affected by vane endwall contouring and shock-boundary 1 ayer interactions on the blade. The program was conducted to ensure timely interaction with the high-pressure turbine component effort, as shown in Figure 1.

T o satisfy the objectives, two vane cascades and three blade cascades were designed and tested. The vane cascades comprised an S-wall cascade, which incorporated the same endwall shape envisioned f ~ r the component vane outer diameter platform, and a straight-wall cascade, which served as a baseline configuration for comparison. The blade cascades comprised three different airfoil geometries: (1) base, ( 2 ) overcambered, and (3) straightback. The base airfoi 1 represented the component design; the overcambered airfoi 1 featured more canber toward the trailing edge than the base design, and the straight- back design featured a f l a t t e r suction surface downstream of the throat than the base design.

This report presents the program t e s t procedures and results associated with the testing of these cascades. I t i s divided into two major sections; Section 3.0, which discusses the vane cascade program and Section 4.0, which discusses the blade cascade program. Section 5.0 contains the overall program conclusions.

ORIGINAL PAGE 19 OF POOR QUALITY HIGH PRESSURE TURBINE COMPONENT DESIGN AND FABRICATION SUPERSONIC CASCADE TEST PROGRAM 'D - CRITICAL DECISION POINT FOR SELECTION OF VANE & BLADE AERODYNAMICS - COMPONENT PRELIMINARY DESIGN COMPLETED 2 - COMPONENT DETAILED DESIGNED INITIATED 3 - COMPONENT DETAILED DESIGN COMPLETE3 4 - AIRFOIL SELECTION CRITERIA PROVIDED TO TEST PROGRAM FROM COMPONENT PRELIMINARY DESIGN EFFORTS 5 - BENEFITS OF S-WALL CONFIRMED, BASE BLADE AIRFOIL SHAPE CONFIRMED AS PROVIDING BEST PERFORMANCE 6 - VANE AND BLADE LOSS CHARACTERISTICS WITH COOLING FLOW VERIFIED F i g u r e 1 Supersonic Cascade Program Logic Diagram

ORIGINAL PAGE fl

OF POOR QUALITY 3.0 VANE CASCADE PROGRAM 3.1 Analysis and Design Published 1ite:ature along w i t h P r a t t & Whitney A i r c r a f t cascade and engine t e s t r e s u l t s have shown t h a t t u r b i n e i n l e t vane losses can be reduced b y contouring the outer diameter p l a t f o r m i n t h e manner i l l u s t r a t e d i n Figure 2.

This contouring reduces t h e vane i n l e t Mach nuntier and reduces t h e d i f f u s i o n on the suction s i d e o f the vane near the t r a i 1 i n g edge; b o t h o f which can impfpye t u r b i n e p e r f o r n e through a r e d u c t i o n i n secondary losses. (Deich e t .

a 1 and Ewen et. a 1 T 2 f ) . Both r e p o r t achieving increases i n e f f i c i e n c y i n r o t a t i n g r i g s a t t r i b u t e d t o enawall p r o f i l i n g .

INNER DIAMETER

/ PUTFORM

CYLINDRICAL FLOW ENDWALL

-I

ENDWALL

.\

OUTER DIAMETER PLATFORM Figure 2 Schematic o f Turbine I n l e t Vane Endwall Contour research i n the area o f endwall p r o f i 1 i n g was conducted b y M o r r i s and PrevioY3) Hoare , who i n v e s t i g a t e d a l i n e a r cascade f i t t e d w i t h several d i f f e r e n t endwall geometries. This t e s t i n g was conducted a t an e x i t Mach number i n t h e incompressible range and a t a R ~ y n o l d s number about an order o f magnitude below t h a t o f t y p i c a l commercial engine s e r v i c e conditions. Also, e x i t a i r angle data were not obtained so t h a t mass averaged losses could not be p r o p e r l y determined a t the measurement plane.

Since the Energy E f f i c i e n t Engine high-pressure t u r b i n e component i n l e t guide vane i s a high t u r n i n g design employing a contoured endwall, as w e l l as i n t e r - nal c o ~ i d u c t i v e and external f i l m cooling, i t was d e s i r a b l e t o f 1) v e r i f y t h e expected performance b e n e f i t associated w i t h t h e contoured endwall and ( 2 ) determine the performance penalty associated w i t h the i n j e c t i o n o f coolant f l o w i n t o "he f l o w f i e l d surrounding the a i r f o i 1.

ORIGlNAL PAGE 13

OF POOR QUALITV An a d d i t i o n a l o b j e c t i v e was t o employ measured data t o assess t h e accuracy of a n a l y t i c a l methods and t o g a i n a physical understandin o f t h e effect of end- ! w a l l p r o f i l i n g on the vane cascade intra-passage f l o w i e l d . These methods range from secondary l o s s c o r r e l a t i o n s t o t h e numerical modeling o f the three- dimensional i n v i s c i d f l o w f i e l d .

To s a t i s f y t h i s objective, two vane cascades were designed: (1) an S-wall cascade, which incorporated t h e same endwall shape envisioned f o r t h e high- pressure t u r b i n e component vane outer diameter p l a t f o r m and (2) a s t r a i g h t w a l l cascade, which served as a base1 i n e c o n f i g u r a t i o n f o r comparison. Flow- paths for these cascades are compared i n Figure 3. To conduct the c o o l i n g f l o w tests, t h e a i r f l o w i n t h e s t r a i g h t w a l l cascade was designed t o incorporate i n t e r n a l flow passages (see Figure 7 ( b ) ) .

S-WALL CONFIGURATION GAGING LINE Figure 3 S t r a i c ~ h t Wall and S-Wall Cascade Configurations

ORIGINAL PAGE I S

OF POOR QUALITY The aspect ratio of each configuration, based on axial chord, equaled the 1.4 aspect ratio of the corresponding vanes for the high-pressure turbine component. The t i p section airfoil geometry from a 43 percent reaction design was selected because the effects of the flow characteristics at the vane air- f o i l outer diameter platform iniersection were of primary interest. The 43 percent react ion level was chosen for the cascades becau5e component design studies had shown this high reaction to provide slightly better performance than a lower reaction design. Both cascades used this same untwisted spanwise section geometry, as shown in Figure 4. Refer t o Appendix A for a listing of the airfoil coordinates used for the S-wall cascade (Table A-1) and the straight wall cascade (Table A - 2 ) .

43 PERCENT REACTION TIP SECTION STRAIGHT S-WALL

AXIAL CHORD - CM (IN) 3.340 11.315) 3465 (1.364)

PITCH- CM (IN) 8.466 (3.333) 8.466 (3.333) THROAT- CM IINJ 1.425 10.561 1.516 (0.597) LEADING EDGE RADIUS- CM IlnJ) 0.523 (0.206) 0.523 10.206) TRAILING EDGE RADIUS- CM (IN) 0.053 10.021 0 053 (0.021) INLET METAL ANGLE !DEG) 90.00' INLET WEDGE ANGLE iDEG) 9O.OO0 EXIT METAL ANGLE (DEG) 10.43' EXIT WEDGE ANGLE IDEG) 4.00' UNCOVERED TURNING'(DEGI 1 2.0O0

ACTUAL CHORD - CM (IN) 9.012 (3.548) 9.012 (3.548

'NOTE: THE STRAIGHT WALL VANE IS ROTATED OPEN 0.67' Figure 4 Vane Cascade Geometry ORlGlNAL PAGE i3 OF POOR QIJALITY Mean1i n e a n a l y s i s was performed t o o p t i m i z e t h e t u r b i n e c o n f i g u r a t i o n . Primary c o n s i d e r a t i o n s d u r i r ~ g t h e mean1 i n e a n a l y s i s were i n l e t and e x i t Mach t r i a n g l e s and gas t u r n i n g angles. A s t r e a m l i n e a n a l y s i s was subsequently employed t o generate r a d i a l p r o f i l e s . This a n a l y s i s u t i 1 i z e d b o t h twoand three-dimensional a n a l y t i c a l procedures t o produce s u i t a b l e a i r f o i 1 pressure d i s t r i b u t i o n s . The a n a l y s i s a l s o determined t h e r a d i a l d i s t r i b u t i o n o f aerodynamic p r o p e r t i e s .

Boundary l a y e r c a l c u l a t i o n s were then used t o v e r i f y t h e l o s s c h a r a c t e r i s t i c s o f t h e a i r f o i l s . Through t b i s approach, c r i t e r i a were e s t a b l i s h e d f o r t h e design o f b o t h cascades.

The vane a i r f o i ! sections were designed so t h a t t h e f l o w was accelerated p a s t t h e gage p o i n t ( t h r o a t ) w i t h low, smooth backend d i f f u s i o n . The uncovered t u r n i n g and e x i t wedge angle were optimized t o minimize t h e two-dimensional loss. An i t e r a t i v e procedure using a computer i n t e r a c t i v e a i r f o i 1 design system was used t o design t h e e x t e r n a l contours o f t h e a i r f o i l s and t o estab- 1 i sh t h e d e s i r e d a i r f o i 1 s t a t i c pressure d i s t r i b u t i o n .

The p r e d i c t e d pressure d i s t r i b u t i o n f o r sach cascade a t 50-percent span and t h e pressure d i s t r i b u t i o n a t t h e vane t i p s e c t i o n o f t h e high-pressure t u r b i n e are presented i n F i g u r e 5. The vanes o f t h e s t r a i g h t component ( f o r reference) w a l l cascade were r o t a t e d open t o p r o v i d e t h e cascade w i t b an e x i t a i r angle equal t o t h e S-wall cascade. This approach p e r m i t t e d a one-to-one comparison o f t h e e x i t f l o w c o n d i t i o n s . Each cascade was l i m i t e d t o t h r e e a i r f o i l s t o ensure proper f i t i n t h e cascade tunnel. Flow u n i f o r m i t y f r o m a i r f o i l passage t o passage ( p e r i o d i c i t y ) was achieved w i t h these t h r e e a i r f o i l s b y contouring t h e s i dewalls t o match t h e a p p l i c a b l e flow stream1 i n e as e s t a b l i s h e d b y p o t e n t i a l f l o w computations. I n a d d i t i o n , t h e s i d e w a l l s were a d j u s t a b l e t o p e r m i t minor p e r i o d i c i t y adjustments once t h e cascade was i n s t a l l e d i n t h e tunnel (see F i g u r e 6).

- ENGINE(HPT COMPONENT)

0-0- S-WALLCASCADE CASCADE AERODYNAMICS

- - STRAIGHT WALL CASCADE

MACH MACH AIR AIR MAX NO. NO. ANGLE ANGLE MACH IN OUT IN OUT NO.

-----

ENGINE (HPT COMPONENT) .070 .84 90.00° 10.43O .98 S-WALL CASCADE ,082 .84 90.00° 10.43O .95 STRAIGHT WALL CASCADE .I00 .84 90.00° 10.43O .92 F i g u r e 5 Compariscn o f Pressure D i s t r i b u t i o n s ORIGINAL PP.GE :3 OF POOR QLIALIW 1. AIRFOIL SECTION - 43% REACTION VANE TIP SECTION 2. SIDEWALLS ARE STREAMLINES FROM AERODYNAMIC ANALYSIS INDICATES COOLANT DISCHARGE LOCATIONS ADJUSTABLE Figure 6 Cascade Configuration To understand how cooling a i r a f f e c t s performance, t h e f l o w r a t e f o r each c o o l i n g s i t e must be accurately known. When one plenum feeds several c o o l i n g s i t e s , i t i s not always possible t o c o n t r o l the c b o l i n g f l o w s p l i t t o w i t h i n desired l i m i t s . Therefore, i n t h i s cascade design each c o o l i n g s i t e was metered by a separate plenum. Two changes t o the i n t e r n a l c o o l i n g f l o w passage scheme envisioned f o r t h e high-pressure t u r b i n e component vane, shown i n Figure 7(a) were necessary t o execute t h i s approach: (1) t h e showerhead cool ing holes were eliminated, and ( 2 ) t h e pressure s i d e f i l m cool i n g was "conbined" a t one i n j e c t i o n s i t e instead o f being separately metered (see Figure 7 ( b ) ) . The reason f o r the f i r s t change i s t h a t it was mechanically impossible t o p r o p e r l y meter b o t h the showerhead and the s u c t i o n s i d e c o o l i n g f i l m , Since showerhead losses are considered unimportant r e l a t i v e t o suction side losses, showerhead holes were eliminated. The pressure s i d e f i l m c o o l i n g was conbined because i t s r e a r s e t o f f i l m holes were too near t h e t r a i l i n g edge t o permit each s i t e t o be independently metered. D e t a i l s o f t h e cooling geometry are presented i n Table 1.

COMPONENT (a) TEST (b) NOTE: COOLING FLOWS GIVEN AS PERCENTAGE OF PRIMARY FLOW F i g u r e 7 High-Pressure Turbine Blade C o o l i n g Model TABLE 1 V A N E COOLING GEOMETRY No. Hol e Row Locat i o n X/BX Holes D i a . cm ( i n . )

7- 7 m 8 36

2 S S 0.042 3 6 :::! [:::El

SS 0.052 36 0.06 (0.025) 4 PS 0.52 25 0.08 (0.032) 5 PS 0.60 2 5 0.08 (0.032) 6 T E 1 .OO 40 0.05 (0.021)

SS - s u c t i o n s u r f a c e

PS - pressure surface

TE - t r a i l i n g edge

3.3 Testina 3.3.1 - General D e s c r i p t i o n - The vane cascade t e s t program examined two important aspects of t h e vane component design: (1) e r ~ d w a l l c o n f i g u r a t i o n and ( 2 ) c o o l i n g f lcw arrange- ment. I n t h e endwall evaluation, t h e s t r a i g h t w a l l and t h e S-wall cascades were t e s t e d t o compare t h e r e l a t i v e performance o f these endwall con- f i g u r a t i o n s . The cool i n g flow t e s t s were conducted t o assess t h e t o t a l pressure l o s s p e n a l t y associated w i t h coo'l i n g a i r i n t r o d u c t iott ( s e p a r a t e and co&ined) a t t h e pressure and s u c t i o n surfaces and a t t h e t r a i l i n g edge o f t h e vane a i r f o i 1s.

Test F a c i 1 i t y and I n s t r u m e n t a t i o n 3.3.2.1 Test F d c i l i t y -- The P r a t t & Whitney A i - c r a f t Plane Case- "o Wind Tunnel ( T e s t Stand X-32) i s a steady f l o w tunnel c o n s i s t i n g a f a -ge plenum, t e s t s e c t i o n , and discharge c e l l . F i g u r e 9 presents a s c h e m a ~ i c r e p r e s e n t a t i o n o f t h i s f a c i l i t y . The primary a i r f l o w e n t e r s t h e plenum chamber and i s discharged against t h e plenum endwall. The f l o w subsequently passes through a honey- c o d f l o w s t r a i g h t e n e r and f i n e mesh screens, which remove s w i r l and make t h e f l o w uniform b e f o r e i t e n t e r s a r e c t a n g u l a r be1 lmouth t o t h e cascade approach duct. A f t e r passing through t h e cascade, t h e a i r discharges t o t h e t e s t c e l l , which i s maintained a t atmospheric pressure. Cascat- i n c i - dence a i r angles a r e s e t b y r o t a t i n g t h e cascade assembly r e l a t i v e t o t h e d i r e c t i o n o f t h e a i r f l o w i n t h e approach duct.

A square b a r g r i d was placed i n t h e duct downstream o f t h e b e l l m o u t h en- t r a n c e i n o r d e r t o increase t h e t u r b u l e n c e i n t h e f i o w , thus m i n i m i z i n g any p o t e n t i a l f l o w s e p a r a t i o n w i t h i n t h e cascade. For t h i s c o n f i g u r a t i o n , t h e p r e d i c t e d f r e e strearn t c r b u l e n c e l e v e l ( u l / u ) was 2.7 p e r c e n t a t t h e entrance t o t h e cascade 17 inches downstream o f t h e g r i d . Hot w i r e measurements were taken a t t h i s l o c a t i o n , w i t h o u t a span reducing s i d e - p l a t e and a t a Mach number o f 0.1, corresponding t o t h a t encountered i n t h e cascad2 t e s t i n g . The measured value o f 2.6 percent u l / u was i n ex- c e l l e n t agreement w i t h t h e p r e d i c t i o n . I n s t a l l a t i o n o f span r e d u c i n g s i d e p l a t e s employed i n t h e t e s t i n g produced an area c o n t r a c t i o n r a t i o of approximately 1.5 downstream o f t h e g r i d , causing a r e d u c t i o n of f r e e stream turbulence i n t o t h e cascade. This c o n t r a c t i o n was ca!cul ated t o reduce +he t u r b u l e n c e l e v e l f r o m 2.6 t o 1.8 percent.

Instrunerstat i o n -- The vane cascade was equipped w i t h endwall s t a t i c pressure t a p s t o o b t a i n d a t a on endwall cross-passage pressure g r a d i e n t s (see F i g u r e 10). Because t h e s t r a i g h t w a l l cascades were symmetrical about t h e 50 p e r c e n t span l o c a t i o n , o n l y one o f t h e endwalls r e q u i r e d s t a t i c pressure taps, The asymmetrical S-wall cascade, however, r e q u i r e d taps on b o t h endwalls. The center a i r f o i l o f t h e s t r a i g h t w a l l cascade was p r o v i d e d w i t h pressure taps as shown i n F i g u r e 11. The a i r f o i l s on each s i d e were p r o v i d e d w i t h pressure taps on t h e s i d e b o r d e r i n g a f u l l passage and a t t h e t r a i l i n g edge. The center a i r f o i l o f t h e S-wall cascade was equipped w i t h s t a t i c taps as shown i n F i g u r e 12. As w i t h t h e s t r a i g h t w a l l cascade, a i r f o i l s vane passage.

on e i t h e r s i d e had s t a t i c taps on t h e s i d e b o r d e r i n g a f u l l PRIMARY AIRFLOW ';9ure 9 High-Pressure Turbine Vane Cascade Test F a c i l i t y BOTH ENDWALLS FOR S-WALL CASCADE ONE ENDWALL FOR STRAIGHT WALL CASCADE Figure 10 High-Pressure Turbine Vane Cascade Endwall S t a t i c Pressure Tap Locations The instrumentation employed f o r vane cascade t e s t i n g i s presented i n Table 2.

Both the cone probe and t h e cobra probe were r u n i n a f r e e j e t c a l i b r a t i o n f a c i 1 i t y t o develop c a l i b r a t i o n curves f o r t o t a l p .essure, s t a t i c pressure, yaw angle, and p i t c h angle (5-port conbination probe o n l y ) . This c a l i b r a t i o n was conducted a t approximately the same u n i t Reynolds number as the cascade e x i t f l o w and over the range o f Mach numbers and angles r e q u i r e d f o r t h e t e s t i n g .

TABLE 2 VANE TEST PROGRAM INSTRUMENTATION

I-ocat i o n Measurement Type Q u a n t i t y

Tunnel Plenum T o t a l Temperature Thermocouple 1 T o t a l Pressure K i e l Probe 1 Approach Duct S t a t i c Pressure S t a t i c Taps 3 Discharge C e l l Barometric Pressure Barometer I Coolant Supply Flow Rate Rot ameter I (each i n j e c t i o n s i t e ) T o t a l Pressure P i t o t Probe 1 T o t a l Temperature Thermocouple 1 Survey Plane T 3 t a l Pressure Cone Probe (1) 1 Downstream o f S t a t i c Pressure Cob! 7 Probe ( 2 ) 1 Cascade Yaw Angle P i t c h Angle Vane Surf aces S t a t i c Pressure S t a t i c Taps ( S t r a i g h t Endwal i Cascade) Vane Endwa 1 1 s S t a t i c Pressure S t a t i c Taps ( S t r a i g h t Endwall Cascade) Vane Surf acss S t a t i c Pressure S t a t i c Taps 5 5 (S-Wall Cascade) Van? Endwall s S t a t i c Pressure S t a t i c Taps 30 ( S-Wal 1 Cascdde) (1) The cone probe i s a 5 - p o r t c c h i n a t i o n probe used t o o b t a i n measurements o f t o t a l pressure, s t a t i c pressure, and p i t c t l n d yaw angles over m s t of t h e t r a v e r s e plane. This probe has a stem diameter o f 3.97 mm 10.156 i n . ) and a c o n i c a l t i p w i t h a 70-degree i n c l u d e d angle.

( 2 ) The cobra probe c o n s i s t s o f t h r e e c a p i l l a r y tubes brazed i n p a r a l l e l . It was used t o neasure f l o w c o n d i t i o n s c l o s e t o t h e endwalls !i .e., w i t h i n t n e boundary 1 ayer) .

EACH STATIC PRESSURE TAP SHOWN ABOVE (0

EACH STATIC PRESSURE TAPShOWN ABOVE (0 1

APPEARS IN EACH SECTION SHOWN BFLOWI--- APPEARS IN EACH SECTION SHOWN BELOW(--\ /"""' " " " 'A ---'------*-- 2 5 1 S P A N

-- - r ----- #-- 50% SPAN

---.------ -- 75% SPAN

AT TRAILING EDGE, PARALLEL TO

I / / / / / / / ///I// /z7

ENDWALL F i gure 12 H i gh-Pressure F i g u r e 11 H i @-Pressure Turbine S-Wall Vane Turbine S t r a i g h t Cascade S t a t i c Wall Vat? Cascade Pressure Tap S t a t i c Pressure Tap L o c a t i o n s Locdt i ons 3.3.3 -- Test Procedures 3.3.3.1 E s t a b l i s h i n g Test C o n d i t i o n s Test c o n d i t i o n s were e s t a b l i s h e d t o p r o v i d e nominal e x i t Mach numbers equal t o t h e design p o i n t e x i t Mach numbers o f 0.84 and 0.92 f o r t h e component vane t i p and mean sections, r e s p e c t i v e l y . These Mach numb r s were o b t a i n e d b y s e t t i n g e x i t s t a t i c - t o - i n l e t pressure r a t i o s o f 0.63 (Mn 0.84) and 0.58 IMn 0.92). The main stream t o t a l t e m p e r a t l ~ r e was a nominal 6 5 O ~ ( 1 5 0 ~ ~ ) . Since t h e f l o w e x i t e d t o atmospheric pressure, t h e t o t a l pressure approaching t h e cascade was i n t h e range o f 20,685 t o 89,635 pa ( 9 t o 13 p s i g ) . The expansion r t i o s of

if

0.63 and 0.58 were c a l c u l a t e d t o y i e l d Reynolds n u h e r s o f 6.0 x 10- , and 7.0 x lo5, based on e x i t f l o w c o n d i t i o n s and on a i r f o i l a x i a l chord. Desigq p o i n t c o o l a n t f l o w r a t e s a r e summarized i n Table 3 (see a l s o F i g u r e 74).

TABLE 3 GESIGN POINT COOLANT FLOW RATES Locat i o n Coolant Flow Rate ( p e r c e n t mainstream f l ~ w l S u c t i o n S u r f ace Pressure Surface Tr a i 1 i ng Edge Endwall boundary l a y e r c h a r a c t e r i s t i c s e n t e r i n g t h e cascade were determined by ~neasurements taken a t a l o c a t i o n i n t h e approach d u c t 2.54 cm ( 1 i n 1 i n f r o n t o f t h e l e a d i n g edge plane o f t h e cascade. These data were necessary t o f u l l y c h a r a c t e r i z e t h e c o n d i t i o n s a t t h e i n l e t t o t h e cascade. Data were o b t a i n e d f o r t h e range o f Mach nunbers and Reynolds numbers t o he encountered d u r i n g t e s t i n g . A cobrz probe was used t o o b t a i n these data. I n t e g r a l boundary l a v e r parameters were determined f o r t h e measbred v e l o c i t y p r o f i l e s and were found t o be i n reasonable agreement w i t h those c a l c u l a t e d l ~ s i n g a well-accepted f o r m u l a t i o n f o r t h e development o f a zero pressure g r a d i e n t f u l l y t u r b u l e o t boundary l a y e r o r i g i n a t i n g a t t h e g r i d . I n t e r p o l a t i o n between t h e measured d a t a was subsequently used t o o b t a i n t h e i n l e t displacement and momentum thicknesses f o r t h e two cascades a t t h e i r r e s p e c t i v e t e s t p o i n t c o n d i t i o n s .

These a r e g i v e n i n Table 4. For t h e endwall comparison, f u l l passage e x i t plane surveys and s u r f a c e s t a t i c pressure data were o b t a i n e d f o r each cascade a t a nominal e x i t i s e n t r o p i c Mach number o f 0.85. Table 5 presents t h e measured t e s t c o n d i t i o n s f o r each cascade. These c o n d i t i o n s a r e r e p r e s e n t a t i v e o f a l t i t u d e c r u i s e c o n d i t i o n s f o r t h e high-pressure t u r b i n e component design.

TABLE 4 INLET BOUNDARY LAYER PARAMETERS S-Wall S t r a i g h t Wall Cascade Cascade Displacement Thicknsss 0.117 cm (0.046 i n . ) 0.124 cm (0.049 i n . ) Momentum Thickness 0.089 cm (0.035 i n . ) 0.092 cm (0.036 i n . ) Momentum Thickness 2900 Reynolds No.

Displacement 0.0130 Thickness/Chord TABLE 5 TEST CONDITIONS S t r a i g h t Wall S-wall Cascade Cascade I s e n t r o p i c E x i t Mach NP. 0.849 0.845 Upstream A i r Angle ( a l ) 90 degrees 90 degrees Upstream Mach No. 0.109 0.091

Reynolds No. (Re) B X 6.41 x l o 5 7.00 x l o 5

Upstream T o t a l 160,653pa(23.29PSIA) 162,032pa(23.53PSIA) Pressure T o t a l Temperature 624O2 (347OK) 574OR (3lg0K) 3.3.3.2 Shakedown T e s t i c g Shakedown t e s t i n g c o n s i s t e d o f p r e s s u r e leak checks and c a l i b r a t i o n o f a l l i n s t r u m e n t a t i o n b e f o r e performance t e s t i n g was i n i t i a t e d . A p r e l i m i n a r y data p o i n t was r u n t o v e r i f y performance o f t h e i n s t r u m e n t a t i o n and d a t a a c q u i s i - t i o n systems. The performance t e s t program was i n i t i a t e d a f t e r i t was ascer- t a i n e d t h a t a l l i n s t r u m e n t a t i o n and systems were o p e r a t i n g p r o p e r l y .

Performance T e s t i n g The vane cascade t e s t program was s t r u c t u r e d t o p e r m i t separate performance e v a l u a t i o n s o f ( 1 ) endwall c o n f i g u r a t i o n s ( S - w a l l . and s t r a i g h t w a l l cascades) and ( 2 ) c o o l i n g f l o w e f f e c t s ( i .e., performance e f f e c t s of t h e t h r e e areas of c o o l i n g f l o w i n j e c t ion: t r a i 1 i n g edge, s u c t i o n surface, and pressure surface).

The t e s t s conducted t o evaluate performance focused on t o t a l pressure loss, a i r f o i 1 pressure d i s t r i b u t ian, and e x i t angle.

Wake t r a v e r s e data were used t o assess cascade performance i n terms o f t o t a l pressure loss. These t r a v e r s e s were made 1.02 crn (0.4 i n . ) downstream o f t h e t r a i l i n g edge. The 5 - p o r t c o b i n a t i o n probe was used t o o b t a i n measurements of t o t a l pressure, s t a t i c pressure, and p i t c h and yaw angles over most o f t h e t r a v e r s e plane. This probe was t r a v e r s e d i n t h e p i t c h w i s e d i r e c t i o n a t a constant span h e i g h t t a k i n g measurements a t 0.15 cm (0.060 in.) increments.

Yaw angles (angles i n plane para1 l e l t o endwalls) were o b t a i n e d b y nu1 1 i n g t h e probe aerodynamically t o w i t h i n one degree and then employing c a l i b r a t i o n curves. The probe d r i v e a x i s o f r o t a t i o n passed through t h e t i p o f t h e probe.

P i t c h angles were o b t a i n e d f o r t h e 5 - p o r t c o m i n a t i o n probe through t h e c a l i b r a t i o n curves. Each f u l l passage e x i t survey n o m i n a l l y c o n s i s t e d o f 35 p i t c h w f s e t r a v e r s e s covering t h e f u l l span o f t h e cascade.

For t h e c o o l a n t i n j e c t i o n t e s t s , a i r was metered t o t h e vane plenums. Coolant t o t a l pressure, t o t a l temperature, and f l o w r a t e were measured.

3.3.4 Performance Test Plan The t e s t p l a n f o r t h e vane cascade t e s t s i s shown i.1 Table 6. T h i s p l a n was formulated t o meet t h e f o l l o w i n g t e s t o b j e c t i v e s : Estciblish t h e performance i n terms o f t o t a l p r e s s u r e l o s s and e x i t angle o f o r t h e S-wall c o n f i g u r a t i o n r e l a t i v e t o t h e s t r a i g h t w a l l cascade, and o E s t a b l i s h t h e t o t a l p r e s s u r e l o s s p e n a l t y associated w i t h c o o l a n t i n j e c t i o n a t d i f f e r e n t l o c a t i o n s i n t h e vane ( s e p a r a t e and c o r b i n e d i n j e c t i o n ) .

Data Reduction and A n a l y s i s The d a t a a c q u i s i t i o n sequence f o r t h e vanes i s presenteo i n Table 7.

The d a t a a n a l y s i s methods f o r t h e vanes are: 1. Comparison of t h e s t a t i c pressure d i s t r i b u t i o n between t h e S-wall and s t r a i g h t w a l l c o n f i g u r a t i o n s .

2. Comparison o f measured a i r f o i 1 s t a t i c pressures w i t h a n a l y t i c a l p r e d i c t ions.

3. Comparison-of t h e f u l l passage pertormance, i n terms o f t o t a l pressure l o s s and e x i t a i r angle, between t h e S-wall and s t r a i g h t w a l l con- f i g u r a t i o n s .

4. Comparison o f t o t a l pressure losses w i t h a n a l y t i c a l p r e d i c t i o n s ( w i t h and w i t h o u t c o o l ant e j e c t i o n ) .

5. Determination o f t h e performance s e n s i t i v i t y t o c o o l i n g f l o w r a t e v a r i a t i o n s .

6. Determination o f t h e performance s e n s i t i v i t y t o e x i t Mach n u h e r v a r i a t i o n s .

3.3.6 Experimental U n c e r t a i n t y Experimzntal u n c e r t a i n t i e s f o r t h e r e s u l t s obtained a r e estimated t o be t0.02 P/PT for surface s t a t i c data; t0.4 degree f o r gap average mass weighed eSii t a i r angle (a2) and t0.02 for t b e gap average e x i t Mach n u d e r . Mass averaged t o t a l pressure loss-results a r e estimated t o be accurate w i t n i n +5 p e r c e n t t o -8 percent APT/PT i n s i d e the p r o f i l e l o s s r e g i o n and +10 percent t o -14 percentAPT/PT i n s i d e t h e secondary l o s s r e g i o n .

3.4 Results The a n a l y s i s of t h e vane cascade data was d i v i d e d i n t o two areas (1) uncooled cascade t e s t s and ( 2 ) c o o l i n g f l o w discharge t e s t s . The r e s u l t s o f t h i s a n a l y s i s are discussed i n t h e f o l l o w i n g sections.

TABLE 6 VANE CASCADE TEST PLAN Endwall E x i t ~ x ~ a n s i o n ( l ) Test Shape MachNo. R a t i o C o o l i n g Data Scan S t r a i g h t 0.84 0.63 No c o o l a n t e j e c t i o n F u l l passage e x i t plane survey S t r a i g h t 0.88 0.66 No c o o l a n t e j e c t i o n Mid-span e x i t plane sur s'ey 0.92 0.58 0.96 0.55 S t r a i g h t 0.84 0.63 T r a i 1 i ng edge, F u l l ?assage e x i t pressure s i d e and plane survey s u c t i o n s i d e e j e c t i o n a t design p o i n t c o o l a n t f l o w r a t e S t r a i g h t 0.84 0.63 T r a i 1 i n g edge Mid-span e x i t plane e j e c i i o n a t survey design p o i n t c o o l a n t f l o w r a t e S t r a i g h t 0.84 0.63 Pressure s i de Mid-span e x i t e j e c t i o n a t 70, 100, p l a n e survey and 130 percent of design p o i n t c o o l a n t f l o w r a t e S t r a i g h t 0.84 0.63 Suction s i d e Mid-span e x i t e j e c t i o n a t 70, 100, p l a n e survey and 130 percent o f design p o i n t c o o l a n t f l o w r a t e S t r a i g h t 3.92 0.58 S u c t i o n s i d e Mid-span e x i t e j e c t i o n a t 70, 100, p l a n e t r a v e r s e and 130 perceot o f design p o i n t c o o l a n t f l o w r a t e 5-wall 0.84 0.63 No c o o l a n t e j e c t i o n F u l l passage e x i t p l a n e survey R a t i o of E x i t S t a t i c t o I n l e t T o t a l Pressure.

TABLE 7 Sequence Data Obtained 1. Apply known pressures t o Transducer c a l i b r a t i o n transducers 2. Set cascade expansion r a t i o None 3. Set c o o l a n t f l o w r a t e s Primary f l o w i n l e t t o t a l pressure Primary f l o w temperature C e l l s t a t i c pressure A i r f o i l s t a t i c pressures Endwall s t a t i c pressures Coolant f l o w r a t e s Coolant t o t a l pressures 4. Program probe c o n t r o l l e r Flow f i e l d e x i t t r a v e r s e f o r l o c a l and s t a r t data a c q u i s i t i o n t o t a l pressure, s t a t i c pressure, sequence p i t c h angle, and yaw angle 5. Check cascade expansion Check f o r d r i f t o f t e s t r a t i o and c o o l a n t f l o w r a t e s ; c o n d i t i o n s ; t e s t cnded i f t o be repeated p e r i o d i c a l l y s i g n i f i c a n t d r i f t occurs d u r i n g data a c q u i s i t i o n sequence 6. A f t e r t e s t i s completed, Check o f transducer c a l i b r a t ion; r e p e a t (1) r e p e a t t e s t i f c a l i b r a t i o n has d r i f t e d .

Flow V i s u a l i z a t i o n Gefore t h e s t a r t o f performance t e s t i n g , s u r f a c e f l o w v i s u a l i z a t i o n s were conducted f o r b o t h cascades t o assess t h e behavior o f t h e l i m i t i n g s t r e a m l i n e s and t o e s t a b l i s h whether t h e r e were any f l o w s e p a r a t i o n problems. These flow v i s u a l i z a t i o n s were made b y a p p l y i n g a m i x t u r e o f lampblack and o i 1 t o t h e a i r f o i 1 and endwall surfaces and subsequently o p e r a t i n g t h e cascade t u n n e l a t t e s t p o i n t c o n d i t i o n s f o r approximately one minute. F i g u r e 13 presents t h e f l o w v i s u a l i z a t i o n achieved f o r t h e s t r a i g h t w a l l cascade viewed f r o m up- stream. F i g u r e 14 shows t h e S-wall cascade viewed from downstream. Both cascades e x h i b i t e d t h e same qua1 i t a t i v e f e a t u r e s found i n previous t e s t i n g of a l a r g e scale cascade o f t u r b i n e a i r f o i l s .

ORIGINAL PAGE ' @ CF POOR QUALITY The f e a t u r e s o f a three-dimensional s e p a r a t i o n a t t h e cascade i n l e t a r e shown Here, t h e i n l e t boundary l a y e r i s seen t o separate along l i n e s ' n F i g u r e 13.

S1 and S2 t o f o r m a "horseshoe" shaped v o r t e x c o n t a i n i n g t h e low momentum boundary l a y e r f l u i d (see F i g u r e 1 5 ) . The l e g corresponding t o S1 gets wrapped around t h e s u c t i o n s u r f a c e o f t h e a i r f o i l t o form what has Seen c a l l e d t h e "counter vortex," w h i l e t h e l e g corresponding t o S2 moves toward t h e s u c t i o n s u r f a c e o f t h e adjacent a i r f o i 1 t o form t h e passage vortex. The new boundary l a y e r formed w i t h i n t h e r e g i o n bounded b y s e p a r a t i o n l i n e S2 and attachment l i n e s A 1 and A2 i s swept toward t h e s u c t i o n s u r f a c e o f t h e adjacent a i r f o i l b y t h e c r o s s channel pressure g r a d i e n t . T h i s low momentum f l u i d subsequently c o n t r i b u t e s t o t h e grcwth o f t h e passage vortex, making i t t h e dominant f e a t u r e o f cascade secondary f l o w . Toward t h e r e a r o f t h e a i r - f o i l , s e p a r a t i o n l i n e S moves o n t - t h e s h c t i o n s u r f a c e o f t h e adjacent a i r - f o i l as can be seen i n ? i g u r e 14 o f t h e cascade e x i t plane.

From t h e f l o w v i s u a l i z a t i o n s o f t h e r e a r o f t h e s u c t i o n surface, t h e s t r a i g h t w a l l cascade f e a t u r e s were found t o be symmetric about midspan; those o f t h e S-wall cascade were s l i g h t l y asymmetric (see F i g u r e 14), which i s t o be expected. I n p a r t i c u l a r , t h e passage v o r t e x s e p a r a t i o n l i n e a l o n g s i d e o f t h e s t r a i g h t w a l l was found t o b e c l o s e r t o t h e endwall than t h e opposing S-wall s i d e (appr?uimately 5.08 mm (0.20 in.) a t t h e t r a i l i n g edge o f t h e s t r a i g h t w a l l s i d e cc,.ipared t o 6.35 rnm (0.25 i n . ) f o r t h e S-wall s i d e ) . This d i s t a n c e was found t o b e approximately 6.35 mm (0.25 i n . ) f o r b o t h sides o f t h e s t r a i g h t w a l l cascade.

Figure 1 5 Cascade Secondary Flow Schematic 3.4.2 P e r i o d i c i t y Evaluation To achieve a good cascade performance evaluation, i t i s important t h a t t h e f l o w c h a r a c t e r i s t i c s w i t h i n adjacent vane passages (see Figure 16) be as as possible. That i s t o say good p e r i o d i c i t y must be nearly i d e n t i c a l achieved. To assess the passage f l o w c h a r a c t e r i s t i c s , s t a t i c pressure data were obtained on adjacent a i r f o i l s i n each cascade. A i r f o i l - t o - a i r f o i 1 comparison o f these data were used t c j established p e r i o d i c i t y w i t h i n t h e cascade b e f o r e any performance t e s t i n g was conducted. A s can be seen b y t h e pressure d i s t r i b u t i o n data i n Figures 17 and 18 (which a r e r e p r e s e n t a t i v e o f the other spanwise measurements taken), b o t h t h e s t r a i g h t w a l l and t h e S-wall cascades showed excel l e n t agreement between t h e passage readings corresponding t o the adjacent f l o w passages, i n d i c a t i n g t h a t e x c e l l e n t p e r i o d i c i t y was attained.

3.4.3 Pressure D i s t r i b u t i o n Analysis S t a t i c pressure data were also obtained f o r each cascade a t t h e cross-channel l o c a t i o n s i d e n t i f i e d i n Figure 10 and spanwise l o c a t i o n s i d e n t i f i e d i n Figures 11 and 12. (See Section 3.3.2.2, Instrumentationl. These data v e r i f i e d t h e p r e d i c t i o n techniques used t o c a l c u l a t e t h e f l o w w i t h i n the cascade.

Cross-channel s t a t i c pressure measurements obtained a t one endwall of the s t r a i g h t w a l l cascade are shown i n Figure 19 along w i t h the p r e d i c t e d r e s u l t s from a two-dimensional compressible p o t e n t i a l f l o w c a l c u l a t i o n .

The asymmetric geometry o f the S-wall cascade produces a three-dimensional f l o w f i e l d . Consequently, t h e two-dimensional p r e d i c t i o n method employed f o r the S-wall cascade was replaced w i t h a three-dimensional i n v i s c i d f l o w f i e l d This method was used t o p r e d i c t the cross-channel and ca!culation method.

span~tise pressure d i s t r i b u t i o n s i n the S-wall cascade. Cross-channel data were taken on b o t h the s t r a i g h t w a l l and S-wall, and spanwise data were recorded a t the spanwise l o c a t i o n s shown i n Figure 12. Results are presented i n Figures 20 through 23.

The three-dimensional eFfects o f the S-wall are most noticeable by comparing the cross-channel data i n Figures 19 and 20. The e f f e c t o f t h e S-wall on the s t r a i g h t endwall i s t o decrease the maximum Mach number on the s u c t i o n surface and s h i f t i t s l o c a t i o n from approximately 50 percent chord t o 70 percent chord.

These f a c t o r s c o n t r i b u t e t o a reduction i n the endwall cross-channel pressure gradient w i t h a subsequent r e d u c t i o n i n secondary loss. (See s e c t i o n 3.4.4.1.2 f o r a more d e t a i l e d discussion.) Overall, there i s good agreement between the t e s t data and the predictions.

3.4.4 Uncooled Cascade Results Uncooled cascade t e s t data were analyzed w i t h regard t o making a comparison between the s t r a i g h t w a l l design and the S-wall design. I n addition, the performance o f the s t r a i g h t w a l l cascade was assessed i n terms o f e x i t Mach number v a r i a t i o n s .

ORIGINAL ! ' I " . ,.: ' OF POOR QUA!. : ' I - Figure 16 Cascade Flow Passages M p = O . 8 d PASSAGE 1

n PASSAGE 2

-PREDICTION "

F i g u r e 17 S-Wall Cascade Mid-Span Pressure D i s t r i b u t i o n P e r i o d i c i t y Evaluation '4 M2 = 0.83 0 PASSAGE 1 0.4 0 PASSAGE 2

- PREDICTION

t

F i g u r e 18 S t r a i g h t Wall Mid-Span Pressure D i s t r i b u t i o n P e r i o d i c i t y E v a l b a t i o n

-

PREDICTION 96 GAP DATA SUCTION SURFACE

=

--.-

. 75

-0- 100 PRESSURE SURFACE F i g u r e 19 S t r a i g h t - h a l l Cascade - Comparison o f Cross-Channel Data With 2 5 Two-Dimensional P o t e n t i a l Flow Computation ORIGINAL QDGZ L*'?

._ . s .

OF POOR QU.A-I; 4 PREDICTION % GAP DATA

.- - ---

-

i

I - 0 SUCI ,ON SURFACE

----- 25 O

-- 50 0

. 75 0

- - 100 PRESSURE SURFACE

F i g u r e 20 S-Wa1: Cascade - Comparison ~f Cross-Channel D a t a w i t h

Three-Dimensional I n v i s c i d Flow F i e l d Computation a t Strzight W a l l End of Cascade U

I - 0 PLANEA 25%exitspan

F i g u r e 21 S-Wai 1 Cascade - Comparison o f Spanwf s e Data a t Planes A and b w i t h

Three-Dimensional I n v i s c i d Flow F i e l d Cornputat i o n PREDIC-

--

- 0 PLANEC 75%exit span

---- 0 PLANED 86%exitspan

I

F i g u r e 22 S-Wall Cascade - Comparison o f Spanwise Data a t Planes C and l ? w i t h Three-Dimensional I n v i s c i d Flow F i e l d Ccmputation % G A P C ? A

- 0 SUCTION SURFACE

- 5 0

- 100 PRESSURE SURFACE

F i g u r e 23 S-Wall Cascade - Comparison c~f Cross-Channel Data w i t h TI, ?e-Dimensional I n v i s c i d Flow F i e l d Computation a t S-Wall End o f Cascade 3.4.4.1 Comparison o f S-wa I I and S t r a i g h t Wal ! Vane Cascade Results 3.4.4.1.1 Cascade Performance a t Mach 0.84 Cascade performance, judged i n terms o f t o t a l pressure l o s s and e x i t a i r angles, was determined from wake t r a v e r s e data. Contour p l o t s o f t o t a l pressure loss i n t h e e x i t plane were then produced from these data. Figure 24 shows the l o s s contours f o r the s t r a i g h t w a l l cascade over one p i t c h . The a i r f o i l wake i s characterized by near p a r a l l e l contours running spanwise w i t h i n t h e p r o f i l e loss r e g i o n o f span. Nearer t h e endwal l s , t h e r e g u l a r p a t t e r n i s d i s r u p t e d b y the secondary flows. The passage vortex appears i n t h e contour p l o t s as t h e c i r c u l a r r e g i o n o f high-pressure loss near t h e endwalls. Closer t o t h e end- walls, the boundary layers t h a t developed w i t h i n the passage produced a sharp increase i n t o t a l pressure loss.

The pressure l o s s contour p l o t f o r t h e S-wall cascade (see Figure 25) e x h i b i t s features s i m i l a r t o those o f t h e s t r a l g h t w a l l cascade. The o n l y d i f f e r e n c e i s t h a t t h e h i g h l o s s r e g i o n (passage vortex) on t h e s t r a i g h t w a l l r i d e i s smaller and l i e s somewhat c l o s e r t o t h e endwall than t h e h i g h l o s s r e g i o n on the S-wall side. This i s consistent w i t h t h e s u c t i o n surface separation l i n e s observed i n t h e f l o w v i s u a l i z a t i o n s conducted.

Comparison o f the spanwise d i s t r i b u t i o n s o f mass-averaged t o t a l pressure l o s s f o r b o t h cascades (see Figures 26 and 27) shows t h a t t h e most notable feature o f the S-wall cascade i s the apparent lack o f the t o t a l pressure l o s s peak, caused by the vortex on t h e s t r a i g h t w a l l side. The vortex on t h i s s i d e was closer t o the endwall, and p i t c h w i s e i n t e g r a t i o n of l o s s data merged t h i s loss r e g i o n i n t o t h a t o f t h e endwall boundary layer.

Yaw angle data were mass-averaged i n the p i t c h w i s e d i r e c t i o n i n t h e same manner as t h e pressure l o s s data. The yaw angle d i s t r i b u t i o n f o r t h e s t r a i g h t w a l l cascade (Figure 28) shows good symmetry about midspan. The passage v o r t i c e s near the endwalls r e s u l t i n a r e g i o n o f f i r s t underturning and then overturning, proceeding from midspan across the v o r t i c e s toward t h e endwalls.

Closer t o the endwal l s , t h e f l o w again e x h i b i t s decreased turning. This i s a t t r i b u t e d t o the suction surface l e g o f the leading edge counter vortex (see Figure 1 5 ) . The p r e d i c t i o n o f t h e cascade e x i t a i r angle i s presented i n Figure 28 and i s seen t o y i e l d a s l i g h t l y smaller average e x i t a i r angle than measured.

The spanwise d i s t r i b u t i o n o f yaw angle f o r the S-wall cascade (Figure 29) shows the same o v e r a l l t u r n i n g as the s t r a i g h t w a l l cascade. The f l o w on the S-wall side, however, e x h i b i t s a ~ p r e c i a b l y more overturning. This t r e n d i s w e l l predicted, i n an average sense, b y t h e three-dimensional f l o w f i e l d c a l c u l a t i o n shown, which suggests t h a t the i n v i s c i d c o n t r i b u t i o n t o t h e behavior o f t h e endwall f l o w i s appreciable.

3.4.4.1.2 Loss Assessment Total cascade loss can be t r e a t e d as the sum o f the p r o f i l e (two-dimensional] loss associated w i t h the geometric shape o f the a i r f o i l s e c t i o n and t h e secon- dary loss r e s u l t i n g from endwall boundary l a y e r e f f e c t s . Understanding these loss elements i s c r u c i a l t o t h e assessment o f a p a r t i c u l a r design.

40 60 PERCENT S Figure 24 Straight W a l l Cascade Pressure Loss Contours Y

S-WALL 5 J I / , 1 . 0

L

CURVE LABEL

-

2 0.02 4 0.04 6 0.06 8 0 08 1 0 0.10 12 0 12 STRAIGHT

r WALL

40 60 PERCENT S F i g u r e 25 S-Wall Cascade P r e s s u r e Loss Contours ORIGINAL PWG; $2 OF POOR QL'AL1.N PROFILED b- PLANAR WP.! L ' '' PERCENT SPAN F i g u r e 26 Spacwise D i s t r i b u t i o n o f Total P r e s s u r e Loss f o r S-Wall Cascade 0 2 0 40 60 80 100 PERCENT SPAN F i g u r e 27 Spanwise D i s t r i b u t i o n o f T o t a l P r e s s u r e Loss f o r S t r a i g h t Wall Czscade 1 5 a2 10

--- PREDICTION

0 20 40 60 80 100 PERCENT SPAN F i g u r e 28 Spanwise D i s t r i b u t i o n o f E x i t Yaw Angle f o r t h e S t r a i g h t W a l l Cascade a2 1 0 STRAIGHT WALL

---- PREDICTION

PERCENT SPAN F i g u r e 29 Spanwise D i s t r i b u t i o n o f E x i t Yaw Angle for t h e S-Wall Cascade P r o f i l e Loss A n a l y s i s Because o f t h e s t r o n g o v e r a l l a c c e l e r a t i o n experienced b y t h e f l o w i n t h e cascades under i n v e s t i g a t i o n , i t would b e expected t h a t even a t t h e r e l a t i v e l y low aspect r a t i o of 0.5, t h e midspan r e g i o n would b e reasonably two-dimensional i n nature, and t h e r e f o r e amenable t o two-dimensional a n a l y t i c a l methods.

The r e s u l t s f r o m t h e pressure d i s t r i b u t i o n c a l c u l a t i o n s were used t o execute two-dimensional boundary l a y e r computations f o r b o t h t h e pressure and s u c t i o n surfaces. For t h e S-wall cascade, t h e pressure d i s t r i b u t i o n remains n e a r l y constant over 75 percent o f t h e span on t h e S-wall side, j u s t i f y i n g t h e approach. A wake mixing c a l c u l a t i o n was subsequently used t o "mix o u t u * t h e boundary l a y e r s and t o account f o r t h e f i n i t e t h i c k n e s s o f t h e a i r f o i l t r a i 1 i n g edge.

Boundary l a y e r c a l c u l a t i o n s f o r b o t h cascades p r e d i c t e d t h a t t r a n s i t i o n occurs on t h e s u c t i o n s u r f a c e s l i g h t l y downstream o f t h e s t a r t o f t h e adverse pres- sure g r a d i e n t . The pressure s u r f a c e i n b o t h instances was p r e d i c t e d t o remain laminar t o t h e t r a i l i n g edge.

The p r e d i c t e d t o t a l pressure l o s s f o r b o t h cascades u s i n g t h i s procedure was w i t h i n 12 percent o f t h e values o b t a i n e d b y m i x i n g o u t t h e r e s p e c t i v e wake t r a v e r s e data. Table 8 summarizes these r e s u l t s .

TABLE 8 MIDSPAN REGION PRESSURE LOSS COMPARISON P r e d i c t e d Pressure Measured Pressure LOSS ( A P I P ) LOSS ! AP/P) S t r a i g h t Wall Cascade 0.0134 0.0120 S-Wal 1 Cascade 0.0124 0.0118 Secondary Loss A n a l y s i s The wake t r a v e r s e plane was chosen a t a d i s t a n c e downstream o f t h e cascade where t h e f l o w i s s u f f i c i e n t l y mixed o u t t o p e r m i t a good assessment o f t h e secondary losses t o be made. I n a d d i t i o n , t h e t e s t r e s u l t s were analyzed on b o t h an area-averaged and a mass-averaged b a s i s . Secondary l o s s r e s u l t s were obtained b y s u b t r a c t i n s t h e p r o f i l e l o s s from t h e f u l l passage l o s s . P r o f i l e l o s s was evaluated b y using t h e a p p r o p r i a t e massor area-averaged l o s s over 20 percent o f t h e span about midspan. The r e s u l t i n g values f o r secondary l o s s thus i n c l u d e c o n t r i b u t i o n s from t h e i n l e t boundary l a y e r , endwall boundary layer, and t h e i n t e r a c t i o n between t h e endwall and a i r f o i 1 boundary l a y e r s .

Refers t o t h e a n a l y t i c a l mixing o f t h e measured two-dimensional "*Mix Out" wake t r a v e r s e data t o a homogeneous s t a t e through t h e a p p l i c a - t i o n o f t h e equations f o r conservation o f mass and momentum. I t i s a technique commonly used t o c a l c u l a t e t o t a l cascade l o s s .

3 3 Table 5 presents t h e r e s u l t s f o r b o t h t h e S-wall and t h e s t r a i g h t w a l l cascades i n terms o f mass and area-averaged values. These r e s u l t s a r e presented f o r each h a l f span, as w e l l as over t h e f u l l passage, i n o r d e r t o b r i n g o u t t h e e f f e c t o f p r o f i l i n g one endwall. I n general, t h e r e s u l t s i n d i c a t e t h a t t h e secondary l o s s comprises over h a l f o f t h e t o t a l loss. The area-averaged l o s s does n o t account f o r t h e reduced mass f l o w i n r e g i o n s o f higher t o t a l pressure l o s s . Consequently, t h e area-averaged losses a r e somewhat higher t h a n t h e mass-averaged values.

TABLE 9 Secondary Loss Measurements A ) Planar Wall Cascade Mass A v e y a e - Area Averaged % Span 0-50 50-100 0-100 0-50 50-100 0-100 T o t a l APT/PT .022 '025 .023 .027 .032 .030 P r o f i l e APT/PT .011 .011 .011 .012 .012 .012 Secondary APTIPT .O12 .014 .013 .016 .020 .018 B) P r o f i l e d Wall Cascade % Span 0-50 50-100' 0-100 0-50 50-loo* 0-100 T o t a l APT/PT .018 .021 .019 ,021 .030 ,026 P r o f i 1 e APTIPT .010 .010 .010 .012 .032 .012 Secondary APT/PT .007 .011 .009 .010 ,019 .014

* P r o f i l e d w a l l s i d e

Comparison of t h e mass-averaged data f o r t h e two cascades shows t h a t t h e S-wall cascade has 17 p e r c e n t l e s s f u l l passage l o s s t h a n t h e s t r a i g h t w a l l cascade. Since t h e measured " p r o f i l e " losses were approximately equal f o r b o t h cascades, t h i s improvement was i n t h e secondary losses ( approximately 30 p e r - c e n t r e d u c t i o n ) . Comparison o f t h e d a t a f o r each h a l f span i n d i c a t e s t h a t approximately 65 percent o f t h i s secondary l o s s improvement occurred on t h e s t r a i g h t w a l l h a l f o f t h e S-wall cascade. Area averaging o f t h e d a t a a l s o i n secondary l o s s f o r t h e s t r a i g h t w a l l s i d e o f shows a s u b s t a n t i a l r e d u c t i o n t h e S-wall cascade. Area averaging o f t h e secondary losses f o r t h e S-wall h a l f , however, shows n e g l i g i b l e change f r o m t h e r e s u l t s o f t h e s t r a i g h t w a l l cascade t e s t s . The reason f o r t h i s r e s u l t i s t h a t t h e increased t u r n i n g near t h e S-wall causes a r e d u c t i o n i n t h e mass flow through t h i s r e g i o n which i s n o t accounted f o r b y area-averaging.

An understanding o f secondary f l o w s and t h e avai l a b i 1 i t y o f r e 1 i a b l e p r e d i c - t i o n methods i s c r u c i a l f o r t h e desSgn o f low aspect r a t i o c o n f i g u r a t i o n s . The f l o w i n t h e endwall region, however, i s extremely complicated, as evidenced b y previous f l o w v i s u a l i t a t ions. Consequently, t h e p r e d i c t i o n o f cascade sec0ndar.v l o s s i s m o s t l y l i m i t e d t o e m p i r i c a l c o r r e l a t i o n s of experimental data.

M o r r i s and ~ o a r e ( ~ ) o b t a i n e d secondary l o s s data f o r a s t r a i g h t w a l l cascade f e a t u r i n g i n l e t de vanes w i t h 6 5 degrees o f t u r n i n g . T h e i r data, along w i t h

?dl

t h e data of Came f o r t h e same c o n f i g u r a t i o n , a r e presented i n F i g u r e 30 i n terms o f t h e r e s u l t i n g c o r r e l a t i o n . While t h e M o r r i s and Hoare r e s u l t s a r e r e p o r t e d t o be mass-averaged, w i t h o u t l o c a l angles and v e l o c i t i e s b e i n g obtained, t h e d a t a would a c t u a l l y r e f l e c t more o f an area averaging. F o r t h i s reason, t h e c u r r e n t area-averaged s t r a i g h t w a l l cascade d ? i a have been i n c l u d - ed i n F i g u r e 30 and a r e i n good agreement w i t h p r e d i c t i o n s .

A CAME DATA

PLANAR WALL CASCADE DATA

- MORRIS AND YOARE CORRELATION

X = 0.294(6*/c) +0.011 F i g u r e 30 Comparison o f S t r a i g h t Wall Cascade Data w i t n t h e C o r r e l a t i o n o f M o r r i s and Hoare The secondary l o s s r e s u l t s o b t a i n e d f o r t h e S-wall cascade a r e presented i n F i g u r e 31, along w i t h t h e d a t a o f M o r r i s and Hoare f o r t h e i r " P r o f i l e A" con- f i g u r a t i o n , which i s s i m i l a r t o t h e c u r r e n t p r o f i l e contour. Again, area- averaged losses have been r e p o r t e d f o r reasons p r e v i o u s l y discussed. Also shown i n F i g u r e 31 a r e t h e c o r r e l a t i o n s developed f o r t h i s p a r t i c u l a r endwall geometry. As i n d i c a t e d , t h e c u r r e n t d a t a f a l l w i t h i n t h e s c a t t e r band of the M o r r i s and Hoare data.

An e x p l a n a t i o n o f these experimental r e s u l t s can b e s t be made i n terms of t h e endwall pressure d i s t r i b u t i o n ( l o a d i n g ) . I n comparing t h e l o a d i n g of t h e s t r a i g h t w a l l o f t h e S-wall cascade (see F i g u r e 20) w i t h t h e l o a d i n q on end- w a l l s o f t h e s t r a i g h t w a l l cascade (see F i g u r e 19). i t can be seen t h a t ( 1 \ t h e S-wall cascade endwall i s n o t as h e a v i l y loaded, ( i .e., t h e average pressure s u r f a c e t o s u c t i o n s u r f a c e pressure d i f f e r e n t i a l i s reduced) and I ? ) the center of pressure i s l o c a t e d f a r t h e r a f t . These features c o n t r i b u t e t o a r e d u c t i o n i n t h e cross-passage pressure g r a d i e n t which, i n t u r n , reduces t h e er>dwall cross-passage boundary l a y e r f l o w (see F i g u r e 15). I t i s t h i s low mo~nentum boundary l a y e r f l u i d which c o n t r i b u t e s t o t h e growC$ o f t h e passage vortex and a corresponding increase i n secondary losses. ,,ugh a d e t a i l e d e v a l u a t i o n o f t h e endwall boundary l a y e r c h a r a c t e r i s t i c s , n o t w l t h i n t h e scope o f t h i s program e f f o r t , reduced cross-passage b o u n d a l l l a y e r f l o w i s f e l t t o be t h e p r i m a r y c o n t r i b u t o r t o t h e reduced secondary losses noted.

ORIGINAL PAGE iS V [7 MORRIS AND HOARE PROFILE WALL

0 MORRIS AND HOARE PLANAR WALL

PROFILED WALL DATA PLANAR WALL DATA

- MOPRIS AND HOARE PROFILE WALL CORRELATION

- - - MORRIS AND HOARE PLANAR WALL CORRELATION

F i g u r e 3 1 Comparison o f S-Wall Cascade Data w i t h t h e ' P r o f i l e A ' C o r r e l a t i o n o f M o r r i s and Hoare A comparison between t h e loadings o f t h e p r o f i l e d endwall o f t h e S-wall cascade (see F i g u r e 23) and t h e s t r a i g h t w a l l cascade (see F i g u r e 19) a l s o shows t h e p r o f i l e d endwall i s n o t as h e a v i l y loaded and i t s center o f pressure i s f u r t h e r a f t ; again suggesting improved performance, which i s c o n s i s t e n t w i t h t h e data. However, t h e secondary l o s s r e d u c t i o n a t t h e S-wall i s o n l y about one-half t h a t achieved a t t h e s t r a i g h t w a l l ; even w i t h t h e reduced load- ing. Two p o s s i b l e reasons f o r t h i s are (1) t h e maximt~m Mach numbers f o r t h e two endwalls a r e approximately equal and ( 2 ) t h e 5-wall has a s l i g h t l y l a r g e r wetted surface.

3.4.4.2 E f f e c t s o f E x i t Mach N u h e r V a r i a t i o n on S t r a i g h t Wal! Cascade Performance The s t r a i g h t w a l l vane cascade was designed u s i n g t h e high-pressure t u r b i n e component vane t i p s e c t i o n a i r f o i l geometry and i n c o r p o r a t e d no t w i s t i n t h e spanwise d i r e c t i o n . I n t h e ccir~ponent vane design, a s l i g h t amount o f t w i s t i s i n c o r p o r a t e d and t h i s changes t h e f l o w passage c h a r a c t e r i s t i c s between a d j a - cent a i r f o i l s such t h a t e x i t Mach nunber increases from t i p t o r o o t (i.e., 0.84 and 0.92 f o r t h e component vane t i p and r o o t s e c t ions, r e s p e c t i v e l y ) . I t was desirable t o examine t h i s spanwise Mach number v a r i a t i o n i n order t o determine i f transonic drag r i s e (evidenced b y a sharp increase i n pressure loss) might occur near the r o o t section. Since the a i r f o i l geometry o f the r o o t , mean, and t i p sections i s n e a r l y i d e n t i c a l , changing cascade e x i t Mach n u h e r provided the desired simulation.

The r e s u l t s o f these t e s t s are shown i n the predicted and measured pressure loss data o f Figure 32 and i n the predicted and measured pressure d i s t r i b u - t i o n s shown i n Figures 33 through 36. These r e s u l t s i n d i c a t e good agreement betweeri predicted and meas~~red data. More importantly, they i n d i c a t e t h a t t h e r e was no abrupt pressure r i s e , even though the e x i t Mach number approached sonic conditions, Flow i n the r o o t area o f the component vane, therefore, can b e expected t o be f r e e o f transonic drag r i s e , confirming the aerodynamic acceptab i 1 i t y o f the vane design.

0 DATA

- PREDICTION

- I

i DESIGN POINT MACH NO.

Figure 32 S t r a i g h t Wall Cascade Midspan Pressure Loss vs. Mach Number E x i t a i r angle v a r i a t i o n w i t h Mach number i s compared t o a n a l y t i c a l p r e d i c t ions i n Figure 37. F a i l u r e i n the t e s t equipment precluded measurements a t other than the t i p section design p o i n t Mach number, however, data taken a t a higher Mach number i n t h e suction surface i n j e c t i o n t e s t s indicated, as predicted, t h a t e x i t a i r angle i s r e l a t i v e l y i n s e n s i t i v e t o e x i t Mach number v a r i a t i o n (see Figure 5 5 ) . Good agreement i s shown between p r e d i c t i o n and t e s t data at the design p o i n t Mach n u h e r .

ORIGINAL PACE : S

OF POOR QlJAi-I'lf 0 DATA

- PREDICTION

I

F i g u r e 33 S t r a i g h t Wall Cascade P r e d i c t e d and Measured Pressure D i s t r i b u t i o n s

- Mach N u h e r = 0.837

3.4.5 Cooled Cascade Results f o r S t r a i g h t Wall Cascade Using t h e s t r a i g h t w a l l cascade, secondary a i r f l o w was discharged from t h e vane t r a i i i n g edge and from t h e pressure and s u c t i o n sides o f t h e a i r f o i l t o e s t a b l i s h t h e e f f e c t s o f these flows on cascade pressure l o s s and e x i t a i r angle. The performance s e n s i t i v i t y t o c o o l i n g f l o w v a r i a t i o n s was assessed a t each l o c a t i o n i n d i v i d u a l l y . I n one t e s t , e x i t Mach number was v a r i e d t o assess t h e cooled cascade c o n f i g u r a t i o n f o r p o s s i b l e drag r i s e a t near sonic Mach numbers. I n another t e s t , a l l c o o l i n g a i r discharge s i t e s were p r o v i d e d w i t h metered f l o w simultaneously t o e s t a b l i s h whether any i n t e r a c t i o n e f f e c t s e x i s t e d . $1 i g h t d i f f e r e n c e s i n e x i t Mach number, Reynolds number, and design p o i n t c o o l a n t f l o w s r e s u l t from t h e i n a b i l i t y t o e x a c t l y d u p l i c a t e t e s t condi- t i o n s f o r each t e s t s e r i e s . The impact o f these s l i g h t d i f f e r e n c e s on t e s t r e s u l t s i s considered t o be n e g l i g i b l e . Actual t e s t c o n d i t i o n s a r e summarized i n Table 10.

' > ORIGINAL I->4,. , , ., OF POOR QSALiTY Figure 34 Straight Wall Cascade Predicted and Measured Pressure Distributions

- Mach Nu&er = 0.877

Figure 35 Straight Wall Cascade Predicted and Measured Pressure Distributions

- Mach Number = 0.906 3 9

3 5 EXIT AIR ANGLE (a) MIXED OUT g 3.

J I P

- DEGREES

Z rt c

3-5

ID -

1 -

TABLE 10 CGOLANf FLOW TEST CONDITIONS E x i t Reynolds W /Wp

Mach No. Number 7%) ToC/Top

A1 1 F i l m Cool i n g S i t e s Flowing (Design P o i n t ) 0.82 6 . 7 ~ 1 0 5 0.96 0.984 T r a i l i n g Edge Pressure S u r f ace 1.27 0.984 Suction Surface 2.49 0.984 Tra i 1 i ng Edge I n j e c t i o n Only 0.83 7.0x105 0.91 0.985 Pressure Surface I n j e c t i o n Only 0.83 6 . 7 ~ 1 0 5 0.87 0.987 0.83 6.7~105 1.23 0.987 0.83 6.7~105 1.50 0.987 Suction Surface I n j e c t i o n Only 0.83 6.8~105 1.71 0.987 (Base1 i rle) 0.83 6.8~105 2.38 0.987 0.83 6 . 8 ~ 1 0 5 3.18 0.987 Suction Surface 0.90 7.7x105 1.43 1.056 I n j e c t i o n o n l y 0.90 7.7~105 2.16 0.986 (Increased 3.90 7 . 7 ~ 1 0 ~ 2.86 0.978 E x i t Mach Nr,) = R a t i o o f c o a l a n t f l o w t o cascade passage i n l e t flow Wc/Wp Toc/Top = R a t i o o f c o ~ l a n t f l o w t o t a l temperature t o cascade passage i n l e t f l o w t o t a l temperature.

To preclude endwall e f f e c t s on t h e i n d i v i d u a l t r a i l i n g edge, pressure side, and suctior! s i d e f l o w i n j e c t i o n t e s t s , cascade performance measurements were obtained o n l y between 25 ana 75 percent span. I n t h i s region, t h e e x i t f l o w c h a r a c t e r i s t i c s were shown t o be n e a r l j two-dimensional i n nature, thus making t h e measurements amenable t o comparison w i t h two-dimensional p r e d i c t i o n methods. For t h e t e s t p o i n t h a v i n ~ a l l t i l e i n j e c t i o n s i t e s f lowin5 a t design p o i n t r a t e s , a f b l l span survey was obtained.

3.4.5.1 E f f e c t s of Simultaneous Design P o i n t Coolant Flow I n j e c t i o n a ; T r a i l i n a ' Edqe and on Pressure and Suction Surfaces o f A i r f o i l Cooling a i r was simultaneously metered t o t h e a i r f o i l t r d i l i n g edge, pressure and s u c t i o n surface s i t e s a t design p o i n i f l o ~ r a t e s . S t a t i c pressure data were subsequently obtained a t 25, 50, and 75 percent span l o c a t i o n s and com- pared t o t h e p r e d i c t e d two-diqensional pressure d i s t r i b u t i o n . The r e s u l t s a r e shown i n F i g u r e 38. This f i g u r e shows t h a t t h e measured vane loading does not vary s i c n i f i c a n t l y i n t h e spanwise d i r e c t i o n and t h a t t h e data a r e w e l l predicted.

ORIGINAL P;l-?.:.'

3 F P X ? R C COOLING FLOW.

- - - P.S. - 1.27 S.S. - 2.49 T.E. - 0.96 '% OF PRIMARY FLOW Figure 38 Pressure D i s t r i b u t i o n vs Chord a t Three Span Locations F u l l passage t o t a l pressure l o s s and a i r angle measurements were obtained a t t h e same l o c a t i o n (1.02 cm (0.4 in.) a x i a l l y downstream o f t r a i l i n g edge plane) as f o r the uncooled cascade. These r e s u l t s are compared w i t h t h e uncooled cascade r e s u l t s i n Figures 39 and 40, and show t h a t t h e e x i t a i r angle wds r e l a t i v e l y i n s e n s i t i v e t o coolant f l o w i n j e c t i o n ; however, t o t a l pressure loss almost doubled i n the two-dimensional f l o w r e g i o n o f t h e span.

The measured e ~ i t f l o w conditions i n t h e midspan vegion (25-75 percent span) Here subsequently mixed-out a n a l y t i c a l l y t o a homogeneous state. This approach permitted a comparison t c be made w i t h a n a l y t i c a l p r e d i c t i o n s . Table 11 compares t h e mixed-out losses w i t h p r e d i c t i o n s f o r t h e uncooled and ccoled cascades. This t a b l e shows t h a t (1) pressure loss increased 75 percent as a r e s u i t o f coolant mixing, and ( 2 ) the predicted cooled pressure l o s s i s higher than the measured value, i n d i c a t i n g t h a t the p r e d i c t i o n method i s conservative.

The f o l l o w i n g sections o f t h i s r e p o r t address the impact on performance associated w i t h f l o w i n j e c t i o n from t h e i n d i v i d u a l s i t e s .

ORIGINAL PAS; t3 OF POOR QilfiLlTJ 0.12

0 NO CO9LING FLOW

DESIGN POlNT COOLING FLOW AT TRAILING EDGE.

0.10 SUCTION SURFACE, AND PRESSURE SURFACE 0.08 ( $ ) 2 06 0.04 0.02 PERCENT SPAN Pressure Loss f o r t h e S t r a i g h t Wall F i g u r e 39 Spanwise D i s t r i b u t i o n o f T o t a l Cascade 3 KO COOLING FLOW DESIGN POINT COOLING FLOW AT TRAILING EDGE.

15 SUCTION SURFACE. AND PRESSURE SURFACE 0 20 40 60 80 PERCENT SPAN F i g u r e 40 Spanwise ? i s t r i b u t i o n o f E x i t Yaw Angle f o r t h e S t r a i g h t Wall Cascade TABLE 11 COMBINED INJECTION PRESSURE LOSS COMPARI%N (Design P o i n t Coolant Flow Rates) Predicted Pressure Measured Pressure Loss ( AP/P) Loss ( 3P/P) C o ~ l e d Cascade Uncoo 1 ed Cascade Cooling Penalty 3.4.5.2 T r a i l i n a Edae Coo:ant Flow Rate E f f e c t s To determine t h e e f f e c t s o f t h e r a t e o f c o o l a c t f l o w i n j e c t i o n a t t h e t r a i l i n g edge, flow was metered t o t h e t r a i l i n g edge s l o t s a t the design p o i n t coolant flow r a t e . Pressure loss and gap-averaged e x i t a i r angle measurements were obtained and compared t o those f o r t h e urlcooled cascade. Test r e s u l t s i n d i - cated t h a t (1) pressure l o s s was increased and ( 2 ) e x i t a i r angle was insensi- t i v e t o coolant f l o w i n j e c t i o n (see Figures 41 and 42).

I n executing a p r e d i c t i o n t o compare w i t h t h e measured cascade performance, the t o t a l pressure o f the coolant f l o w a t t h e t r a i l i n g edge e x j t plane must be known. For a s h o r t coolant f ;ow passage between t h e plenum and t r a i l i n g edge e x i t , it could be assumed t h a t c o o l a r t flow t o t a l pressure a t the t r a i l i n g edge i s equal t o plenum pressure. However, the t r a i 1 i n g edge cooiant f l o w passage i n the high-pressure t u r b i n e component vane design i s long enough so t h a t a s i g n i f i c a n t pressure drop occurs. A methodology was t h e r e f o r e estab- i i s h e d t o account f o r t h i s .

Since plenum and t r a i l i n g edge e x i t s t a t i c pressures and passage flow area were e a s i l y determined, t h e o n l y necessary i n g r e d i e n t remaining t o c a l c u l a t e t r a i l i n g edge e x i t plane t o t a l pressure was coolant f l o w r a t e through the passage. This was estab: ished experimentally b y m t e r i n g the f l o w t o the plenum w h i l e holding t r a i l i n g edge s t a t i c pressure constant and recording plenum pressure a t each f l o w r a t e . The r e s u l t a n t " c a l i b r a t i o n " curve i s shown i n Figure 43. The predicted t h e o r e t i c a l f l o w r a t e shown i n t h e f i g u r e was determined by assuming p i p e f l o w i n t h e passage, along w i t h t h e known passage f l o w area, a ~ i d plenum and t r a i l i n g 3dge e x i t s t a t i c pressures. The r e s u l t s i n d i c a t e e x c e l l e n t agreement between predicted and measwed *;slues o f trz!!in?

edge coolant f;ow over the range o f t e s t program plenum pressures.

"Mixed-out* pressure losses were subsequently c a l c u l a t e d and compared b o t h t o p r e d i c t i ~ n s and t o the uncooled cascade loss. These comparisons showed t h a t t h e measured c o o l i n g penalty i s small and i s we1 1 predicted (see Table 12).

Because o f these r e s u l t s , it was considered unnecessary t o t e s t a t other than design p o i n t c o o l i n g f l o w r a t e s .

-

0.04

0 UNCOOLED VANE

-

0.03

A COOLED VANE

0, n I V) t n

-

9 0.02

W LL I I 25 50 75 PERCENT SPAN F i g u r e 41 Pressure Loss Effects Due t o T r a i l i n g Edge I n j e c t i o n a t Design P o i n t Coolant Flow

0 UNCOOLED VANE

'" A a,,,,,,",..

PERCENT SPAN F i g u r e 42 I i i t A i r Angle E f f e c t s Due t o T r a i l i n g Edge I n j e c t i o n a t Design P o i n t Coolant Flow

ORIGINAL PAGE 19

OF POOR QUALlW TABLE 12 T R A I L I N G EDGE DISCHARGE PRESSURE L O S S COMPARISON ( A p p r o x i m a t e D e s i g n P o i n t ) P r e d i c t e d P r e s s u r e M e a s u r e d P r e s s u r e L o s s ( . \ P / P ) L o s s ( . l P / P ) C o o l e d C a s c a d e U n c o o 1 e d C a s c a d e C o o l i n g P e n a l t y

-

( 1 FOR TWENTY CENTRAL HOLES 0 05 CM (0.02 IN) DIA x 2.08 CM t0.82 1Ni LENGTH (2) FLOW DISCHARGED TO ATMOSPHERIC PRESSURE

-

-

-

-

- PREDICTION

- I I 1 - 0 10 20 30 40 50

PLENUM PRESSURE - INCHES-GAGE

F i g u r e 43 V a n e C a s c a d e T r a i l i n g E d g e F l o w C a l i b r a t i o n 3.4.5.3 Pressure Side Coolant Flow Rate E f f e c t s To assess t h e e f f e c t s o f pressure s i d e c o o l a n t f l o w discharge, f l o w was meter- ed f r o m t h e middle plenum through two rows o f holes l o c a t e d a t t h e b e g i n n i n g o f t h e s t r o n g f a v o r a b l e pressure g r a d i e n t on t h e pressure s u r f a c e o f t h e a i r - f o i l (see F i g u r e 7(b), l o c a t i o n s @ and @ ) . Pressure l o s s and gap-averaged e x i t a i r angle measurements were obtained a t t h r e e c o o l i n g f l o w r a t e s and a r e compared t o those f o r t h e uncooled cascade i n F i g u r e s 44 and 45, r e s p e c t i v e l y .

The s l i g h t skew o f t h e spanwire pressure l o s s d i s t r i b u t i o n s w i t h c o o l a n t f l o w when compared t o t h e uncooled d i s t r i b u t i o n i s b e l i e v e d t o be a t t r i b u t a b l e t o a v a r i a t i o n i n c o o l i n g h o l e f l o w c o e f f i c i e n t s .

"Mixed o u t " pressure losses were subsequently c a l c u l a t e d w i t h t h e spanwise average values b e i n g presented as a f u n c t i o n o f c o o l a n t f l o w r a t e i n F i g u r e 46 along w i t h a p r e d i c t i o n o f t h e data. These pressure losses a t t h e design p o i n t f l o w r a t e a r e compared w i t h p r e d i c t i o n s and w i t h t h e uncooled cascade losses i n Table 13. F i g u r e 47 presents t h e spanwise average e x i t a i r angle as a func- t i o n o f c o o l a n t f l o w r a t e along w i t h a p r e d i c t i o n .

These r e s u l t s i n d ' r a t e t h a t t h e pressure l o s s p e n a l t y f o r pressure s u r f a c e c o o l a n t discharge i s small, t h e l e v e l b e i n g reasonably w e l l p r e d i c t e d . E x i t a i r angle r e s u l t s i n d i c a t e t h a t t h e angle i s i n s e n s i t i v e t o c o o l a n t f l o w r a t e v a r i a t i o n s and i s w e l l p r e d i c t e d .

TABLE 13 PRESSURE SURFACE DISCHARGE PRESSURE LOSS COMPARISON (Approx. Design P o i n t ) P r e d i c t e d Pressure "Mixed Out" Pressure Loss ( .\P/P) Loss ( .\P/P) Coo 1 ed Cascade 0.0144 0.0139 Uncoo 1 ed Cascade Cool i n g Penalty 0.0010 0.0019 3.4.5.4 Suction Side Coolant Flow Rate E f f e c t s To assess t h e e f f e c t s o f s u c t i o n s i d e c o o l a n t f l o w discharge, f l o w was metered from t h e forward plenum through t h r e e rows o f holes l o c a t e d i n t h e a c c e l e r a t i n g f l o w r e g i o n o f t h e a i r f o i 1 s u c t i o n surface (see F i g u r e 7(b), l o c a t i o n s a, 0,

and 0 ) . Pressure l o s s and gdpaveraged e x i t a i r angle measurements were ob-

t a f ~ e d a t t h r e e c o o l i n g f l o w r a t e s a t a 0.84 e x i t Mach number. These r e s u l t s a r e compared t o those f o r tile uncooled cascade i n F i g u r e s 48 and 49.

ORIGINAL PAGE OF POOR QUALIIk

-

0.04

0 NO COOLANT FLOW

A DESIGN POlNT COOLANT FLOW

'

0.03 0 71 % DESIGN POINT COOLANT FLOW

a, a )( 122% DESIGN POINT COOLANT FLOW Q tn tn

-

J 0.02 LU a V) tn W a PERCENT SPAN F i g u r e 44 Pressure Loss Effects Due t o Pressure S u r f a c e Coolant I n j e c t i o n NO COOLANT FLOW DESIGN POlNT COOLANT FLOW 71 % DESIGN POlNT CO9LANT FLOW 122% DESIGN POINT C9OLANT FLOW 0 1 I 2 5 50 7 5 PERCENT SPAN F i g u r e 45 E x i t Air Angle E f f e c t s Due t o Pressure Surface C o o l a n t I n j e c t i o n

* DATA

- PREDICTION

F i g u r e 46 Pressure Loss E f f e c t s Due t o Pressure Surface Coolant Flow I n j e c t i o n Rate V a r i a t i o n

0 DATA

- PREDICTION

WclWp - PERCENT

F i g u r e 47 E x i t A i r Angle E f f e c t s Due t o Pressure Surface Coolant Flow I n j e c t i o n Rate V a r i a t i o n ORIGINAL PAGE . ; OF POOF! ()!)AL;'F->

-

0.04

0 NO COOLANT INJECTION

A DESIGN POlNT COOLANT INJECTION

0 72% DESIGN POlNT COOLANT INJECTION

X 134% DESIGN POlNT COOLANT INJECTION

-

0.03

e

-4 t 0 I 1 25 5 0 7 5 PERCENT SPAN F i g u r e 48 Pressure Loss E f f e c t s Due t o Suction Surface Coolant I n j e c t i o n

0 NO COOLANT INJECTION

A DESIGN POlNT COOLANT INJECTION

0 72% DESIGN POlNT COOLANT INJECTION

X 134% DESIGN POlNT COOLANT INJECTION

0 l 1

I 2 5 5 0 75 PERCENT SPAN F i g u r e 49 E x i t A i r Angle E f f e c t s Due t o S u c t i o n Surface Coolant I n j e c t i o n ORIGINAL PAGE k 2 OF POOR QUALITY "Mixed o u t " pressure losses were subsequently c a l c u l a t e d w i t h t h e spanwise average values being presented as a f u n c t i o n o f c o o l i n g f l o w r a t e i n F i g u r e 50 along w i t h a p r e d i c t i o n o f the data. These r e s u l t s a t t h e design p o i n t flow r a t e are compared w i t h p r e d i c t i o n s and t h e measured uncooled cdscade losses i n Table 14. As the t a b l e indicates, s u c t i o n surface i n j e c t i o n i s a major c o n t r i - b u t o r t o c o o l i n g losses f o r the vane. Figure 51 presents t h e spanwise average e x i t a i r angle as a f u n c t i o n o f coolant f l o w r a t e along w i t h a p r e d i c t i o n .

40, ) e . - a a Cn M2 = 0.83 DATA

PREDICTIO' -

= 6.8 x 105 1, I 1 1 2 3 WcIWp, PERCENT Figure 50 Vane Suction Surface I n j e c t i o n Midspan Loss vs Cooling Flow Rate TABLE 14 SUCTION SURFACE DISCHARGE PRESSURE LOSS COMPARISON (Apprx. De:,ign Point) Predicted Pressure "Mixed-Out" Pressure Loss ( A P I P ) Loss ( AP/P) Cooled Cascade Uncoo 1 ed Cascade Cool i n g Penalty ORlGlNAL PAG ' :.

1 2 - OF POP? @ I J A I ,'. I

P 1 1 -

0 0 0

I

- V)

2 2

- a i o - W

0 DATA M p = 0.83

gi

a

- PREDICTION ToclTop = 0.987

-

f

k 1 I 1

?$ 8

1 2 3 4 WcIWp, PERCENT Figure 51 S t r a i g h t Wall Cascade E x i t A i r Angle vs Cooling Flow Rate Thtse r e s u l t s c o n f i r m t h e well-known and expected predominance o f s u c t i o n sur- face coolant discharge penalty i n the o v e r a l l vane c o o l i n g loss penalty. This i s best explained by r e f e r r i n g t o F'gure 7 ( b ) . Here i t can be seen t h a t the coolant f l o w i n j e c t e d from the t r a i l i n g edge and pressure surface passages i s more n e a r l y a1 igned w i t h t h e free-stream f l o w around t h e a i r f o i l whereas t h a t i n j e c t e d from the suction surface passages i s n e a r l y normal t o the free-stream flow. I n a d d i t i o n , twice as much coolant f l o w i s i n j e c t e d from t h e suction surface p?ssages as From e i t h e r o f the other two s i t e s . The p r e d i c t i o n was found t o considerably overestimate the measured loss l e v e l . This r e s u l t i s a t t r i b u t e d t o the f a c t t h a t the angle t h a t the c o o l i n g holes make w i t h the suction surface i s approximately 5 5 degrees, a c c n f i g u r a t i o n o u t s i d e t h e data base used i n the a n a l y t i c a l model. It should be noted, however, t h a t t h e l o s s trends w i t h changes i n coolant flow r a t e are w e l l predicted ( i .e., compare t h e slopes o f the two curves). E x i t a i r angle r e s u l t s showed t h a t the angle i s i n s e n s i t i v e t o coolant f l o w r a t e v a r i a t i o n s and i s we1 1 predicted.

3 . 4 . 5 . 5 Fffects of E x i t Mach Number and Suction Side Coolant Flow Rate V a r i a t i o n s on S t r a i g h t Wa? 7 Cascade Performance As was the case f o r the uncooled cascade, e x i t Mach n u h e r was increased t o simulate f l o w conditions near the vane r o o t i n or der t o detemine if t r a n s o n i c drag r i s e might be a problem w i t h a cooled cascade. The s i . c t i o n - s i d e i n j e c t i o n case was selected f o r t h i s e v a l u a t i o n because vane performance was shown t o b e Pressure l o s s and e x i t a i r angle most s e n s i t i v e t o suction surface i n j e c t i o n .

measurements were obtained a t three c o o l i n g f l o w r a t e s .

"Mi4ed-out" pressure losses were subsequently c a l c u l a t e d from these data w i t h the spanwise average values being presented as a f u n c t i o n o f c o o l i n g flow r a t e (see Figure 5 2 ) .

M2 = 0.90

--0-- DATA ToclTop = 0.99

ReBX = 7.2 X l o 5

- PREDICTION

1 2 3 WclWp, PERCENT F i g u r e 52 S t r a i g h t Wall Cascade Pressure Loss vs S u c t i o n Surface Coolant Flow Rate a t E x i t Mach Plumber Equal t o 0.9 As was t h e case when e x i t Mach number was 0.83, t:,2 a n a l y t i c a l p r e d i c t i o n c o n s i d e r a b l y overestimates t h e l e v e l o f pressure loss, although t h e change i n pressure l o s s w i t h v a r i a t i o n s i n c o o l a n t f l a w r a t e agrees q u i t e w e l l w i t h t h e experimental data ( i .e., compare t h e c u r v e s l o p ~ s ) .

F i g u r e 53 presents a comparison o f t h e s u c t i o n s u r f a c e i n j e c t i o n r e s u l t s f o r t h e two e x i t Mach numbers tested, along w i t h t h e i r r e s p e c t i v e a n a l y t i c a l pre- d i c t i o n s .

F i g u r e 54 presents a comparison o f t h e Mach number s e n s i t i v i t y o f t h e vaqe s e c t i o n w i t h no c o ~ l a n t f l o w and f l o w a t t h e design f l o w r a t e . This f i g u r e shows t h a t cascade l o s s increase w i t h Mach number* and s u c t i o n s u r f a c e c o o l a n t discharge i s simila: t o t h e r e s u l t s obtained w i t h no c o o l a n t f l o w .

F i g u r e 55 compares t h e measured e x i t a i r angle r e s u l t s f o r t n e two Mach rumbers t e s t e d and shows t h e e x i t angle t o be i n s e n s i t i v e t o b o t h c o o l a n t f l o w r a t e and Mach number. A p r e d i c t i o n o f t h e data i s also presented i n t h i s f i g u r e , and i s i n good agreement w i t h t h e data.

3.4.6 - Summary o f Vane Cascade Results

3.4.6.1 Uncooled Vane Cascade Design-point mass averaged pressure losses f o r t h e s t r a i g h t w a l l and S-wall cascades a r e summarized i n Table 15. As noted i n t h e t a b l e , p r o f i l e losses f o r t h e two cascades are about t h e same whereas the secondary losses o f t h e S-wall cascade a r e c o n s i d e r a b l y lower than f o r t h e s t r a i g h t w a l l cascade. Because of t h i s , t h e S-wall cascade has a t o t 2 1 pressure l o s s 17 percent lower t h a n t h e s t r a i g h t w a l l cascade. This c o n f i r m s t h e p r e d i c t e d b e n e f i t s associated w i t h contoured endwall s i n t u r b i oe vane cascades.

-- MACH NUMBER

-

0.83

--- DATA

- PREOICf ION

WclWp, PERCENT Figure 53 Effects cf Changes i n Suction Surface Coolant Flow Rate and E x i t Mach Norrbcr on Cascade Pressure Loss 0.03 0 N O COOLANT IWJECTION

a DESIGN POINT SUCTtON SURFACE

COOLANT INJECTION 0.02 0.01

L 1 I

0.6 0.9 1 .O EXIT MACH NO.

Figure 54 Straight Wall Cascade Midspan Loss vs Mach Number With and WIthout r e s i g n P o i n t Coolant Flow Inject ion

4 ZXIT MACH NO. = 0.90

0 EXIT MACH NO. = 0.83

- PREDICTED iN = 0.90

W::/Wp, PERCENT Figure 55 Measured E x i t A i r Angle as a Function o f Suction Surface Coolant Flow V a r i a t i o n f o r E x i t Mach N o b e r s o f 0.83 and 0.9 TABLE 15 U;:"C)OLED STRAIGHT W A L L A N D S-WALL VANE CASCADE PRESSURE LOSSES (Mass Averaged a t E x i t M n = 0.84) Tot a1 Prof i l e Secondary LOSS LOSS LOSS ( I P T / F ~ ) ( .\pT/pT) .ipT/pT) Straight Wal; Cascade 0.023 0.011 0.013 S-Wall Cascade 0.019 0.010 0.009 S-Wall Reduct ion % , I P T / P ~ 17.4 9.1 30.8 Increasing the e x i t Mach number i n the s t r a i g h t w a l l cascade increased pressure losses; however, there was no abrupt increase i n pressure l o s s even though t h e flow approached sonic conditions. This indicated t h a t t h e component bane design i s f r e e o f transonic drag r i s e over the range o f intended opersting conditions.

Spanwise v a r i a t i o n s i n vane exi , jaw angle were e s s e n t i a l l y s i m i l a r f o r the fwo cascades except near the endwall s.

3.4.6.2 Cooled Vane Cascade penal t i e s Design-point mass-averaged mid-span pressure losses and coolant f l o f o r t h e s t r a i g h t w a l l cascade w i t h coolant f l o w i n j e c t i o n from, t h e - i r f o i l t r a i 1 i i . g edge, r-essure surface, and s u c t i o n surface are summarized i n r'able 16. The b a s i c p o i n t s t 3 be made from t h i s summary are: ( I ) The a n a l y t i c a l p r e d i c t i o n method c o n s i s t e n t l y overestimates the prsssure losses.

(2) Design p o i n t pressure loss almost doubled when design p i n t coolant discharge ports. The coolant f l o w penalty f l o w was i n j e c t e d from a1 1 alone i s approximately the same as t h e t c t a l p r o f i l e l o s s f o r the uncooled cascade.

(3) Cool ant i n j e c t i o n penal t i e s associated w i t h t r a i 1 i n g edge and ppessure surface i n j e c t i o n are comparab;e.

( 4 ) The dominant loss penalty r e s u l t s , as expected, from s u c t i o n surface coolant flow i n j e c t i o n .

( 5 ) Increasing e x i t Yach n u h e r causes an increase i n t o t a l l o s s and l o s s penalty. This i s shown b y b o t h t h e measured data and t h e a n a l y t i c a l p r e d i c t i o n . This increase i s o n l y moderate i n che e x i t Mach nunher range o f i n t e r e s t and confirms t h a t f l o w w i t h i n t h e cascade i s f r e e o f transonic drag r i s e .

Predicted and measured e x i t a i r angles were i n good agreement and proved t o be i n s e n s i t i v e t c coolant f l o w i n j e c t i o n r a t e changes as w e l l as changes i n e x i t Mach n u h e r .

TABLE 16 MID-SPAN PRESSURE LOSSES F O R STRAIGIiT WALL CASCADE UITH COOLANT F L O W INJECTIOh ("Mixed Oat" Loss a t Design Point Coolant flows) T r a i 1 i n g Pressure Suction Suction A1 1 Surf ace Surf ace Ports Edae Surf ace I n j e c t i o n I n j e c t i o n f l o w i n g I n j e c t i o n I n j e c t i o n E x i t Mach. No. .83 .R3 . 8 3 .90 .83 Measured Loss w i t h Coolant I n j e c t i o n . \ P T / P ~ .0134 .0139 .0195 .0223 .0210 Measured Loss w i t h N o Coo 1 ant I n j e c t i o n \PT/PT .0120 .0120 .0120 .0140 .0120 Measured Coolant Loss Penalty \PT/PT .0014 .0019 .0075 .008 3 .0090 Predicted Coolant Loss

Penalty ~ \ P T / P ~ .0009 .0010 .Dl81 .03 85 . r)? 78

Predicted Total Loss w i t h Cool ant I n j e c t i o n .\PT/PT .0143 .0144 .0315 .0340 .0312 4.0 BLADE CASCADE PROGRAM 4.1 Analysis - and Design Previous in-house studies o f t u r b i n e blade cascades i n d i c a t e d t h a t a i r f o i l t r a i l i n g edge base s t a t i c pressure and cascade t o t a l pressure ;ass could be s i g n i c i c a n t l y a f f e c t e d by the d i s t r i b u t i o n o f the a i r f o i l surface curvature a f t o f the t h r o a t .

The s i n g l e stage t u r b i n e o f the Energy E f f i c i e n t Engine employs advanced aero- dynaqic c ~ ~ c e p t s t h a t achieve a high r a t i o o f wheel speed t o s p e c i f i c work ( v e l o c i t y r a t i o ) , low r a t i o o f through-f low t o wheel t a n g e n t i a l v e l o c i t y

t

(C,/U) and high AN (product o f annulus area and wheel speed squared). The r e s u l t i s a blade a i r f o i l operating a t supersonic e x i t conditions which i s d i f f e r e n t from s t a t e - o f - t h e - a r t 1m1 t i - s t a g e high-pressure t u r b i n e a i r f o i 1s.

The blade design also employs i n t e r n a l conductive and e x t e r n a l f i l m cool ing.

The conbination o f these f a c t o r s made i t h i c h l y deqirahle t o (1) v e r i f y t h a t t h e d i s t r i b u t i o n o f curvature selected fo:- t h e component blade a i r f o i l geometry -would achieve design o b j e c t i v e s and ( 2 j confirm the predicted e f f e c t s on cas- cade performance due t o t r a i 1i n g edge coolant f l o w discharge f i .e., in-houss studies have indicated t h a t t r a i l i n g edge coolant f l o w discharge e f f e c t s pre- dominate over e f f e c t s caused b y leading edge f i l m csolant f l o w discharge?. An a d d i t i o n a l o b j e c t i v e was t o employ meastlred data t o assess the accuracy o f a n a l y t i c a l methods and t o gain a b e t t e r u d e r s t a n d i n g o f t h e f l o w f i e l d w i t h i n the cascade, p a r t i c u l a r l y as it may be a f f e c t e d b y t r a i 1 i n g edge shocks.

To s a t i s f y these objectives. t h r e e blade cascades were designed t o model t h r e e a i r f o i 1 geometries .hosen f o r evaluation: (1) b?se, ( 2 ) overcambered, and (31 straightback. The base design represented the 43-percent r e a c t i o n component blade mean section.

The overcahered a i r f o i l featured more c a d e r toward the t r a i l i n g edge than the base design; and the straightback design featured a f l a t t e r s u c t i o a sur- face downstream o f the t h r o a t than the Lase a i r f o i 1. Cascade geometry design parameters and predicted pressure d i s t r i b u t i o n s f o r these t h r e e blade designs are summarized i n Figures 56 through 58 r e s p e c t i v e l y . Refer t o Appendix A f o r a l i s t i n g of the a i r f o i l coordinates used f o r the base blade (Table A-31, overcahered blade (Table A-4), and the straightback blade (Table A-5).

The base a i r f o i l design was modified t o include i n t e r n a l f l a w passages jsee Figure 59). These passages provided simulated c o o l i n g a i r f l o w f o r t h e t r a i 1 i n g edge flow discharge tests. These f l o w passages simulate the c o o l i n g a i r d i s - charge geometry conceived f o r the high-pressure t u r b i n e component blade.

The same a n a l y t i c a l methods t h a t were used t o design the vane cascades were also employed i n the blade cascade design. An e x i s t i n g e x i t angle d e v i a t i o n system and r a d i a l #work d i s t r i b b t i o n were modified, using p r e v i o u s l y generated in-house t u r b i n e design c o r r e l a t i o n s , and applied t o the blade design. The blade a i r f o i 1 sections were designed t o the same pressure d i s t r i b u t i o n c r i t e r i a a , t h e vane. The blade suction surface curvature downstream o f the t h r o a t was f u r t h e r optimized t o reduce t r a i l i n g edge shock losses.

CIRIGINAL PAC: 1 3 3 F POOR Q!!L;_i';Y (43% REACTION BLADE MEAN SECTION)

1.4 -

0.4 -

-

0.4

-

0.2 0.3 0.0 1 1 1 0.0 C 0.4 0.6 0.8 1.0 0.00 0.20 0.40 0.60 0.80 1.00 AXIAL CHORD (cm/in.) 2.34 (0.922) EXIT METAL ANGLE (DEGREE) 17.07 PtTCH (cmlir..) 3.45 11.359) EXIT WEDGE ANGLE (DEGREE) 2.00 0.940 (0.370) UNCOVEREDTURNING (DEGREE) 6.00 THRQAT lanIin.1 0.155 (0.061 ) INLET GAS ANGLE (DEGREE) 45.15 W N G EDGE RADIUS (crnlir.)

0.053 (0.021 ) EXIT GAS ANGLE (DEGREE) 17.08 TRAILING EDGE RADIUS (cmlin.)

40.15 MkX MACH NO. 1.68 INLET METAL ANGLE (DEGREE) 30.00 EXIT MACH Ftu. 1.22 INLE r WEDGE ANGLE (DEGREE) Figure 56 High-Pressure Turbine Base Blade Design Parameters and Predicted Pressure D i s t r i b u t i o n AXIAL CHORD Icmiin.) 2.34 (0.922) EXIT METAL ANGLE (DEGREE) 15.50 FITCH Icmlm.) 3.45 (1.359) EXIT WEDGE ANGLE (DEGREE) 2.00 THROAT Icmirn.)

0.940 (0.370) UNCOVERED TURNING (DEGREE) 6.0 LEADING EDGE RADIUS (cmlin.) 0.155 (0.061) INLET GAS ANGLE (DEGRFE) 45.15 TRAILING EDGE RADIUS Icmiin.) 0.053 (0.021 1 EXIT GAS ANGLE (DEGREE) 17.08 INLET METAL ANGLE (DEGREE) 40.15 MAX MACH NO. 1.53 INLET WEDGE ANGLE (L)EGREE) 30.00 EXIT MACH NO. 1.22 F i g u r e 57 High-Pressure Turbine Overca&ered Blade Design Parameters and Predicted Pressure D i s t r i b u t i o n ORIGINAL PAI;Y;S OF POOR 9::. -.; AXIAL CHORD (cn.in.) 2.34 (0.922) EAIT METAL ANGLE (DEGREE) 17.82 3.45 (1.359) EXIT WEDGE ANGLE (DEGREE) 2.00 RTCH (cr:.!~n.)

0.940 (0.370) UNCOVEREDTURNING (DEGREE) 2.00 THROAT (cm in.)

LEADING EDGE RADIUS (cmiin.1 0.155 (0.061 INLET GAS ANGLE (DEGREE) 45.15 0.053 (0.021) EXIT GAS ANGLE (DEGREE) 17.08 TRAILING EDGE RADIUS (crn'ln.)

40.15 MAX MACH NO. 1.57 INLET METAL ANGLE (DEGREE) 30.00 EXIT MACH NO. 1.22 INLET WEDGE ANGLE (DEGREE) Figure 58 High-Pressure Turbine Straightback Blade Design Parameters and Predicted Pressure D i s t r i b u t i o n COOLANT FLOW 0.060 c T.

(0.024 1n.1 0.101 cm (0.040 t n . )

- TRAILING

EDGE 0.472 cm (0.186 In.)

0.106 cm 10.042 In.)

t i o n Figure High-Pressure Turbine Base Blade T r a i 1 i n g Edge Coolant Ejec Geometry 4.2 F a b r i c a t i o n and Assenbly sections Three blade cascade assenbl i e s corresponding t o t h e three a i r f o i 1 being i n v e s t i g a t e d were f a b r i c a t e d f o r t h e program. Each cascade assembly consisted o f e i g h t constant s e c t i o n a i r f o i l s w i t h a span o f 10.16 cm ( 4 in.)

and two c i r c u l a r a c r y l i c p l a s t i c endwalls (Schl i e r e n windows). These endwalls were f r e e t o r o t a t e so t h a t i n l e t a i r incidence angle could be varied. The height o f t h e approach duct was also v a r i a b l e sc t h a t t h e w a l l s would approxi- Two o f t h e e i g h t a i r f o i l s mate t h e stagnation streamlines o f the end a i r f o i l s .

were instrumented w i t h s t a t i c pressure taps. Close t o l e r a n c e p i n s a t the ends o f each a i r f o i l secured the a i r f o i l s between t h e endwalls. I n addition, these p i n s served as conduits f o r t h e t r a i l i n g edge discharge a i r and t h e pressure tap leads. For cooled t e s t i n g , t h e center f o u r a i r f o i l s o f the base a i r f o i l cascade were replaced w i t h a i r f o i l s f a b r i c a t e d w i t h i n t e r n a l f l o w passages (see Figure 59) t o permit t r a i l i n g edge e j e c t i o n o f simulated c o o l i n g a i r .

This 3 i r enters t h e plenum chanber through a feed tube, which extends through the cascade endwall ( s c h l i e r e n window). From t h e plenum t h e f l o w passes through a r e s t r i c t o r p l a t e ( p l a t e w i t h holes i n i t ) , which ensures spanwise f l o w u n i f o r m i t y t o the s l o t s t h a t discharge t h e f l o w a t t h e t r a i l i n g edge o f the a i r f o i l . A f u l l y instrumentated blade i s shown i n Figure 60 and Figure 63 shows an assenbled blade cascade i n s t a l l e d i n the t e s t tunnel.

4.3 Testing 4.3.1 General D e s c r i ~ t i o n The o b j e c t i v e o f the blade cascade t e s t program was t o evaluate t h e p e r f o r - mance o f t h e base, overcahered, and straightback blade designs. The perfor- manc. o f each o f the t h r e e a i r f o i l designs was evaluated i n terms o f t o t a l pressure loss, a i r f o i 1 pressure d i s t r i b u t i o n , base pressure c o e f f i c i e n t , shock loss, and e x i t gas angle. Coolant f l o w t e s t s focused on determining the e f f e c t s o f t r a i l i n g edge coolant f l o w i n j e c t i o n on e x i t gas angle, cascade performance, and cascade f l o w f i e l d .

4.3.2 Test Faci 1 i t y and Instrumentation 4.3.2.1 Test F a c i l i t v The United Technologies Research Center (UTRC) v a r i a b l e d e n s i t y supersonic plane cascade wind tunnel, shown schematically i n Figure 61, i s a steady f l o w tunnel c o n s i s t i n g of a t e s t section, an upstream plenum supplied by a com- pressor, a heat exchaager syctem, and a downstream plenum connected t o an exhauster system. The t e s t section i n l e t duct l e n g t h from the plenum i s 91.4 cm (36 in.) long w i t h 2 two-dimensional contractions. The o v e r a l l c o n t r a c t i o n r a t i o i s approximately 370 t o 1.

The i n l e t duct ( a i r f o i l span) i s 10.7 cm ( 4 in.) long, and i t s height can be v a r i e d from about 10.2 cm ( 4 in.) t o 2!.6 cm (8.5 i n . ) as required when t h e cascade i s r o t a t e d t o change incidence angle.

Mach number and Reynolds number can be independently v a r i e d b y s e l e c t i v e c o n t r o l of the compressor and exhauster. Large P l e x i g l a s windows on b o t h sides o f the t e s t s e c t i o c are provided f o r Schlieren photographs o f shock waves.

Figure 62 provides a schematic representation o f the e i g h t blade cascade i n - s t a l l a t i o n . Figure 63 presents a photograph o f t h e tunnel t e s t s e c t i o n w i t h a cascade i n s t a l 1 ed.

A turbulence g r i d was n o t employed f o r b l a d e cascade t e s t i n g because r e s i d u a l turbulence i n t h e cascade i n l e t f l o w , combined w i t h t h e aerodynamic c h a r a c t e r - i s t i c s o f t h e b l a d e design, precluded t h e p o s s i b i 1 i t y o f f l o w s e p a r a t i o n problerns i n t h e cascade d u r i n g t e s t i n g .

The SchlieI.en system p e r m i t t e d s i n g l e frame r o l l f i l m o r 1 6 mm motion p i c t u r e f i l m t o be taken o f t h e b l a d e cascade. I t s viewing screen p e r m i t t e d m o n i t o r i n g o f t h e image t o be recorded, and i t s c o n t r o l s were l o c a t e d o u t s i d e o f t h e t e s t c e l l .

The p a r a b o l i c m i r r o r s were mounted on l a r g e heavy s t o n e and s t e e l p i e r s t o minimize t h e e f f e c t s o f v i b r a t i o n . The l i g h t beam i n t h e system was f o l d e d t o economize on t h e space a v a i l a b l e i n t h e t e s t c e l l . A dual 1 i g h t source, steady and pulsed, p r o v i d e d t h e i l l u m i n a t ~ o n . Exposures ds s h o r t as 1 / 2 second a t r a t e s as h i g h as 4000 exposures per seccnd were made. Both t h e normal opaque k n i f e edge and t h e o p t i o n a l c o l o r k n i f e edge were o p e r a t o r - a d j u s t a b l e a t t h e camera s t a t i o n f o r optimum c u t o f f . F i g u r e 64 presents a schematic r e p r e s e n t a t i o n o f t h e Schl i e r e n system f o r t h e b l a d e cascade.

I I TE,ST SECTION

t

NOTES - A I L MIRORS ARE ADJUSTED MICROMETER SCREWS - MIRRORS ARE MOUNTED ON HEAVY STONE AND STEEL PIERS - KNII-E EDGE IS ADJUSTABLE IN ORIENTATION AND PENETRATION - THE CAMERA SYSTEMS AVAILABLE ARE 16mm MOVIES OR 6cm x 6cm znd MIRROR "SINGLE SHOT" STILLS PARABOLA - MAXIMUM DIAMETER OF OBSERVABLE \ \ I FOCUSSING FIELD = 3 0 5 c n i ( 1 2 1 n ) - WINDOWS IN TEST SECTION ARE CONSTRUCTED OF PLEXIGLASS F i g u r e 64 Schematic o f S c h i i e r e n System The cascade tunnel i s provided w i t h i n s t r u m e n t a t i o n t h a t p e r m i t s s e t t i n g t h e cascade expansion r a t i o and o p e r a t i n g Rpynolds number. This i n s t r u m e n t a t i o n measures t h e upstream plenum t o t a l pressure and temperature and downstream plenum s t a t i c pressure.

The t u n n e l i s a l s o equipped w i t h a probe sys2em used t o o b t a i n flow f i e l d performance measurements downstream o f t h e cascade. This probe system c o n s i s t s of a probe d r i v e mechanism, probe d r i v e c o n t r o l 1 er, pressure transducers, and data r e c o r d i n g i n s t r u n k z t a t ion. The probe d r i v e mechanism moves t h e probe t o programmed p o s i t i o n s on t h e sJrvey plane downstream o f t h e cascade as i n d i c a t - ed by t h e d r i v e c o n t r o l l e r , This c o n t r o l l e r p e r m i t s b o t h t h e s t e p s i z e and t h e dwell t i m e t o be s e t as r e q u i r e d f o r a given t e s t .

The t e s t f a c i l i t y i s equipped t o p r o v i d e high-pressure a i r t o t h e cascade f o r film c o o l i n g purposes. Coolant a i r f l o w r a t e s a r e measured w i t h rotameters.

I n s t r u m e n t a t i o n i s provided t o e s t a b l i s h t h e t o t a l temperature and t o t a l pressure o f t h e f l o w p r o v i d e d t o t h e cascade.

4.3.2.2 I n s t r u m e n t a t i o n The i n s t r u m e n t a t i o n r e q u i r e d f o r b l a d e cascade t e s t i n g i s presented i n Table 17.

TABLE 1 7 BLADE INSTRUMENTATION

Loca t i o n Measuremer t Type Q u a n t i t y

Supply Plenum T o t a l Pressure K i e l Probe 1 T o t a l Temperature Thermocouple 1 Discharge Plenum S t a t i c Pressure S t a t i c Taps 1 Rot ameter 1 Coolant Supply Flow Rate (each a i r f o i 1) T o t a l P r e s s ~ i - e Pressure Tap I T o t a l Temperature Thermocouple 1 Sur*vey i; 1 ane T o t a l Pressure Wedge Probe 1 Do:instream o f Kie! Probe I Cascade S t a t i c Pi essure Wedge Probe Yaw Angle Wedge Probe B l ade Surf ace S t a t i c Pressure S t a t i c Taps "15 Cascade Passage Schl i e r e n Photograph Schl i e r e n 1 System and Camera *Each Blade: two a i r f o i l s instrumented, as shown i n F i g u r e 65 ORIGINAL PAGE 1 g

3F Pnr)R r>l ! A 1 I'ri

F i g u r e 65 High-Pressure Turbine Blade Cascade S t a t i c Pressure Tap Locat i o n s A 1 1 o f t h e performance d a t a were o b t a i n e d b y u s i n g a wedge probe, shown i n F i g u r e 66, which c o n s i s t s o f a 30-degree i n c l u d e d angle wedge mounted on a stem. A p i t o t tube extends from t h e center o f t h e wedge and meets f l u s h !ditb t h e wedge l e i d i n g edge. Pressure p o r t s f o r sensing a i r a n ~ l e a r e l o c a t e d on b o t h s i d e s o f t h e wedge. The taps f o r sensing s t a t i c pressure are l o c a t e d a t t h e a x i a l l o c a t i o n o f t h e angle sensing t a p s and behind rearward f a c i n g steps.

C a l i b r a t i o n o f t h i s probe was accomplished by r e l a t i n g known f l o w c o n d i t i o n s t o probe pressure readings t o y i e l d c a l i b r a t i o n curves f o r t o t a l and s t a t i c pressure and yaw angle. This c a l i b r a t i o n covered a range o f Mach nunbers from 0.6 t o 1 . 4 .

I n a d d i t i o n t o t h e wedge probe, a nineteen-elenlent k i e l head probe was used t o measurz e x i t t o t a l pressure across t h e b l a d e span i n o r d e r t o determine t h e r e g i o t ~ o f two-dimensional f l o w where more p r e c i s e measurements c o u l d be taken w i t h t h e wedge probe. The elements o f t h i s probe covered approximately 75 percent o f t h e b l a d e span, which was s u f f i c i e n t t o p e n e t r a t e t h e r e g i o n where endwall e f f e c t s b e g i n t o predominate (see, f o r example, F i g u r e 76).

ORIGINAL PAGE fg OF POOR QUALITY TWO G 508MM (0 020CM) DIA. HOLES, O N E HOLE EACH SlDE

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(0.035 IN) A TUBE m.0 0 406 M M 10016CM) DIA HOLES, n N E HOLE EACH SlDE Figure 66 Isometric View o f Wedge Prote T i p 4.3.3 Test Procedures 4.3.3.1 ---- E s t a b l i s h i n g Test C o n d i t i o n s -- Test c o n d i t i o n s were e s t a b l i s h e d t o p r o v i d e a c l o s e s i m u l a t i o n o f engine de- s i g n p o i n t a l t i t u d e c r ~ i s e c o n d i t i o n s . Engine design p o i n t values and t h e corresponding range o f r i g t e s t parameters evaluated a r e shown i n Table 18.

TABLE 18 ENGINE DESIGN POINT VALUES AND RIG TEST PARAMETERS Par ame t er Eng i ne

E x i t Mach Number 1.22 0.6 - 1.2

Reyno 1 ds Nurrt>er 6.0 x lo6

5.0 x 10 *

I n l e t A i r Angle (Degrees) 4 5 33 - 58

Coolant Flow (Percent T o t a l ) 2.0 1.0 - 3.0

*

Although t h e f l i g h t c o n d i t i o n Reynolds number was n o t arktievable i n t h e cascade t u n n e l , t h e value shown i s i n a range where pressure l o s s i s r e l a t i v e l y insensitive t o v a r i a t i o n o f Reyno!d< number.

Shakedown T e s t i n g Shakedown t e s t i n g c o n s i s t e d o f pressure leak checks 3 r d c a l i S r a t i o n o f a'! l instrumentation. A p r e l i m i n a r y data p o i n t was rlrn t o v e r i f y perfot-mance o f t h e insti-umentation and data a c q u i s i t i o n systems. The t e s t program was i n i t i d t e d a f t e r i t was a s c e r t a i n e d t h a t a l l i n s t r u m e n t a t i o n and systems were o p e r a t i n g p r o p e r l y .

4.3.3.3 Performance T e s t i n g The b l c d e cascade t e s t program was s t r u c t u r e d t o p e r m i t performance e v a l u a t i o n s over a range o f t e s t c o n d i t i o n s . These i n c l u d e d v a r i a t i o n s i n c o o l a n t f l o w r a t e f o r t h e t e s t s w i t h t r a i l i n g edge c o o l a n t e j e c t i o n . I n a d d i t i o n , t h e f l o w c h a r a c t e r i s t i c s were observed through t h e use o f S c h l i e r e n techniques.

A i r f o i l s e c t i o n pressure performance was measured b y t r a v e r s i n g t h e wedge probe i n 0.762 mm (0.03 i n . ) s t e p s i n t h e p i t c h w i s e d i r c z t i o n a t midspan of the blade cascade. A minicomputer was used t o c o n t r o l t h e probe t r a v e r s e . A t each sampling l o c a t i o n , t o t a l pressure, s t a t i c pressure, and yaw a n g l e ( a n g l e i n p l a n e o f endwall) were medsured b y p n e u m a t i c - t o - e l e c t r i c a l transducers. The transducer o u t p u t corresponding t o these measu;-ements was recorded and m a g n e t i c a l l y stored, and t h e data were l a t e r reduced.

The cooled base a i r f o i l s were s t a t i c t e s t e d b y f l o w i n g a i r through t h e i r t e p - n a l passages t o atmospheric c o n d i t i o n s , thereby e s t a b l i s h i n g t h e t o t a l pres- sure drop between t h e i n t e r n a l plenum and t h e t r a i l i n g edge. D u r i n g cooled t e s t i n g , tt-? c h a r a c t e r i s t i c s o f Lhe c o o l i n g f l o w d t t h e e x i t o f t h e i n t e r n a l passages c o u l d be c a l c u l a t e d and used i n t h e a n a l y t i c a l p r e d i c t i o n or p e r - formance.

For each t e s t p o i n t , t h e i n l e t a i r i n c i d e n c e angle was s e t b y r o t a t i n g t h e c i r c u l a r endwalls as r e q u i r e d . For each incidence s e t t i n g , t h e h e i g h t o f t h e cascade approach duct was a d j u s t e d so t h a t t h e d u w a l l approximately c o r - responded t o t h e s t a g n a t i o n stream1 i n e s o f t h e outermost a i r f o i 1s i n t h e cascade.

The d e s i r e d Mach n u h e r and Reynolds number f o r a t e s t p o i n t were s e t b y a d j u s t i n g cascade tunnel pressuras and t o t a l temperature t o c a l c u l a t e d values.

Once t e s t c o n d i t i o n s were s t a b i l i z e d , t h e automatic d a t a a c q u i s i t i o n sequence was i n i t i a t e d .

Schlieren techniques were used t o o b t a i n s t i l l photographs which i l ' u s t r a t e d t h e cascade shock patterns. I n some t e s t s , lampblack and o i l were a p p l i e d t o a i r f o i l surfaces t o e s t a b l i s h s u r f a c e f l o w p a t t e r n s .

4.3.4 Performance Test P l a n The t e s t p l a n f o r t h e b l a d e cascade t e s t s i s presented i n Table 19.

TABLE 19 ACTUAL BLADE CASCADE TESTS* Test Nach Nu&:? I n l e t A i r Reyno 1 ds - ( I s e n t r o p i c ; Angle (degrees) Number * The nominal c o o l i n g e j e c t i o n r a t e s were 1, 2, and 3 percent f o r t h e base olade cascade.

E x i t s t a t i c - t o - i n l e t t o t a l pressure r a t i o (1) Each zdscade was evaluated over t h e range o f t e s t c o n d i t i o n s shown w i t h no t r a i l i n g edge coolant 2jectSon. I n a u d i t i o n , t h e base b l a d e cascade was t e s t e d w i t h t r a i l i n g edge c o o l a n t f l o w e j e c t i o n r a t e s o f 1, 2, and 3 percent o f t h e t o t a l flow t o determine t h e performance impact o f c o o l a n t f l o w v a r i a t i o n s c t design p o i n t and o f f - d e s i g n c o n d i t i o n s .

Since a c t u a l e x i t Mach number ( M 2 ) i 5 dependeqt upon t h e a c t u a l cascade ;uss, i s e n t r o p i c Mach number (M2.) was used t o s e t t h e t e s t c o n d i t i o n s . The r e i a t i o i ~ s h i p between M 2nd !or t h e carcade, t e s t e d i s shown i n Figures 67, 68, 6 2 1 % 70.

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I I I 0.5 1 .O 1 5 ISENTROPIC EXlT MP.CH NO., M2i F i g u r e 69 R e l a t i o n s h i p Between Actual E x i t Mach Number and I s e n t r o p i c E x i t Mach dumber f o r t h e Straightback A f r f o i l Cascade 1 I I 0.5 1 .O 1.5 ISENTROPIC EXlT MACH NO., M2i F i g u r e 70 R e l a t i o n s h i p Between Actual E x i t Mach Number and I s e n t r o p i c E x i t Mach Number f o r t h e Base A i r f o i l Cascade w i t h T r a i l i n g Edge Coolant Flow Discharge 4 . 3 . 5 Data - Reduction and - A n a l y s i s The data a c q u i s i t i o n sequence f o r t h e blades i n presented i n Table 20.

TABLE 20 BLADE DATA ACQUISITION SEQUENCE Sequence Data Obtained 1. Apply known pressure t o Transducer c a l i b r a t i o n transducers (*) 2 . Set cascade expansion r a t i o None 3. Set coolan: f l o w r a t e s Primary f l o w i n l e t t o t a l pressure Primary f l o w temperature C e l l s t a t i c pressure A i r f o ; 1 s t a t i c pressures Coolant t o t a l pressures Coolant f l o w r a t e s Local t o t a l pressure, s t a t i c 3. Program probe c o n t r o l l e r pressure, and yaw angle f o r and s t a r t data a c q u i s i t i o n szquence t h e mid-span e x i t plane t r a v e r s e o f two p i t c h e s 5. Check cascacle expansion r a t i o Check f o r d r i f t o f t e s t and c o o l a n t f l o w r a t e s ; c o n d i t i o n s ; t e s t ended i f s i g n i f i c a n t d r i f t occurs t o be repsated p e r i o d i c a l l y dur:ng data a c q u i s i t i o n sequence Transducers are p e r i o d i c a l l y c a l i b r a t e d ' l g a i f i s t r e f e r e n c e (*!

pressures d u r i n g data a c q u i s i t i o n .

The data o b t a i n e d from t h e t e s t program were analyzed i n d e t a i 1 t o p e r m i t t h e f i c a t i o n o f design conc?pts and t h e development o f i n f o r m a t i o n u s e f u l i n ver i . ~ t i n g t h e design of t h e higll-pressure t u r b i n e component. Thts data exec anal y s i s i n c l u d e d t h e f o l l c w i n g : comparison o f measured a i r f o i 1 s t a t i c pressure, w i :ti a n a l y t i c a l 1.

p r e d i c t i o n s .

2. Comparisofi o f measured t o t a l pressure losses w i t h a n a l y t i c a l p r e d i c t iofis.

3 .

Comparison o f measured e x i t a i r angle w i t h a n a l y t i c a l predictions.

Cetermination of t h e p e r f o r n ~ a n i e s e n s i t i v i t y t o inciderlce angle 4 .

v a r i a t i o n s .

5. Determination o f performance s e n s i t i v i t y t o c o o l a n t f l o w r a t e s .

Determination o f t h e variation o f base pressure c o e f f i c i e n t w i t h e x i t 6 .

Mach n u h e r .

Determination a f shock wave l o c a t i o n s .

5 .

Pressure l o s s and e x i t a i r angle data f o r each i n d i v i d u a l cascade a r e mass- averagsd. Summary d a t a used i n the a i r f o i 1 performance comparisons a r e an a r i t h m e t i c average o f t h e a i r f o i l 4 and 5 mass-averaged d a t a measured a t each t e s t c o n d i t i c n . This i s done t o p a r t i a l l y account f o r t h e d i f f e r e n c e i n measured data botween a i r f o i l s 4 and 5 which was due t o n o n - p e r i o d i c i t y i n t h e cascades a t t r a n s o n i c and supersonic Mach numbers.

4.3.6 Experimental U n c e r t a i n t y Experimental u n c e r t a i n t y f o r t h e r e s u l t s o b t a i n e d a r e estimated t o be i0.02 P/PT f o r s u r f a c e s t a t i c data; +2.5 degrees f o r gap averaged e x i t a i r angle (P2j and +.01 i s e n t r o p i c e x i t Rach n u d e r ; +8 p e r c e n t .\PT/PT a t sub- s o n i c e x i t Mach nunbers a1.d - t22 percent \ P ? / P ~ a t supersonic e x i t Mach numbers.

4.4 Results A n a l y s i s o f t h e b l a d e cascade data. described i n s e c t i o n s 4.4.3 and 4.4.4 was (1) t h a t associated w i t h t h e uncooled t e s t s d i v i d e d i n t o two major categories: ( s e c t i o n 4.4.3) and ( 2 ) t h a t associated w i t h t r a i l i n g edge c o o l i n g f l o w d i s - charge t e s t s ( s e c t i o n 4.4.4).

Flow V i s u a l i z a t i o n As w i t h t h e vane cascades, s u r f a c e f l o w v i s u a l i z a t i o n s were conducted p r i o r t o 2erformance t e s t i n g t o e s t a b l i s h i f any f l o w s e p a r a t i o n problems e x i s t e d .

These flow v i s u a l i z a t i o n s were made b y a p p l y i n g a m i x t u r e o f lampblack and o i l t o t h e a i r f o i l surfaces and subsequently o p e r a t i n g t h e cascade t u n n e l a t t e s t p o i n t c o n d i t i ~ n s f o r about one minute. I n a d d i t i o n , because a t design p o i n t c o n d i t i o n s p a r t o f t h e f l o w f i e l d i s supersonic, S c h l i e r e n photographs were o b t a i n e d t o d e t a i l t h e shock s t r u c t u r e .

A t y p i c a l a i r f o i l s u c t i o n s u r f a c e f l o w p a t t e r n a t design p o i n t c o n d i t i o n s ( i n l e t gas angle o f 43 degrees and ex15 Mach number o f 1.25) i s shown ir; F i g u r e 71. This p a t t e r n was observed t o b e , j u a n t i t i v e l y and q u a l i t i v e l y simi l a r f o r t h e base, overcamber, and s t r a i g h t b a c k a i r f o i Is.

R e f e r r i n g t o F i g u r e 71, two s e p a r a t i o n bubble r e g i o n s were observed i n t h e one a t X / B X o f approximately 0.6, ~ h i c h was about 0.25 cm f l o w v i s u a l i z a t i o n s ; (0.1 i n . ) wide and which was also fcund t o o ~ c u r f o r b o t h subsonic and s u p e r s o ~ i c e x i t flows, and a second a t X / B X o f ap?roximately 0.7, which was a l s o about 0.25 cm (0.1 i n . ) wide b u t occurred o n l y f o r supersonic e x i t f l o w s .

The more forward s e p a r a t i o n bubble i s a t t r i b u t e d t o a laminar s e p a r a t i o n r e s u l t i n g from t h e s t a r t o f the adverse pressure g r a d i e n t on t h e s u c t i o n surface (see, f o r example, F i g u r e 82). Boundary l a y e r c a l c u l a t i o n s i n d i c a t e d t h a t t h e boundary l a y e r was laminar a t t h i s p o i n t on t h e s u c t i o n surface. The second separation bubble r e g i o n i n d i c a t e d f u r t h e r a f t on t h e s u c t i o n s u r f a c e i s a t t r i b u t o d t o an o b l ique shock-boundary 1 ayer i n t e r a c t ion. This i s Q'il?.*tdL PAGE

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s . ~ g p c t ~ ci.S by t h e ' * - ' . :r! p h o t 9 g r a p h o b t a i n e d f a r design p o i n t c o n d i t i o n s t q g u r e 7 2 : u n r c h shods t i e presslire s t i r f a r e t r 3 i 7 ing edge shock irnpinginq on t k c h d j a c ~ i n t d t r f c l l s ~ r f a c f ? 31 ~pprttxina,. :. X / & Y = ' 3 . 7 . The two reflected S h Q C k S ~ k W d ~ t i ? r i ~ t : t t , , . ','.-,7~. . icjn ~f a smaf 1 s e p a r l r j0.i b u b b l e as r f e p j c t e d jr?

rhla i:hi.*dt t c 81f the shock inpingement r e g i o n shown i n F i q t i r e ?3. A further d1 S C ~ J S ~ 13n and pret3sent 3 t jf3n o f *'*. siql i e r e n photographs g b t a i q e d for t h e tnfe? .?,-' l sectlonc, t e s t e d 13r-e p r e s e n t e d i n the fa'fowing ~ o c t i ~ n ~ of ' h ; ~ r e p o r t .

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2 e r l o d l c i t ; ', i . e . , t ? ~ r " e j i - e t o + i t - ~ ; - ~ f l o w c h a r a c t e r i s t i c 5 are i d e o t { c a t 2 e t ~ e e n b ! n d e passage, 3nn t f ~ e degree t o 'qhich two-4imension3' f l o w was d l r ' + , ' c . ' -'t cascade tnid-s33r-1 were " s e s s e d a t t-, "" ss?iSsonic a n d supep50r?ic 2x1 t Y - r i a .!.e*- conditions. Val i ? . - t m p ~ ~ i l ; o n s Setdeen a n a l y t i c a l predictions and a i r f o : 1 s e c t ion ~ ~ : r f u r n z n c c - e q k i e e t h j t these ~ c o n d i t inns h e mt. T+e 9a;e a;t.frjt \ c a 5 r a d ~ d 3 s ~ s e d f a r t h i s C S S E Z S ~ ~ ~ ~ ~ ~ . For the purpose o f m a k i n q t h i s ~ S S C S S W ~ T t ' i e s t c r t n d i t i r t n - , Ir*? .judged t o I t t v e nesn achieved w h e n i d e n t i c a i tzto-d11nen: a ~ n n 1 f i o w cundi tionc, e x l s t t p r i t n i n t h e mid-span region o f t h e ~ ~ , r l t ~ . a : d i ~ f ~ ~ I 1 P ~ S S B ~ ~ I . o f t h e carcdde. This mean: t h a t a d j a c e n t , a i r f n i 1% , , u * u e q d a , S ' J ~ ~ ~ C F ! ',?ti s p r e s r ! i c e d i r t r i b ~ t ions and a spanwire p ' t r h averaqe \ . f ~ w n j t v t . a r n t 8 ) l i : p * - ~ f ~ ~ i ~ ~ " e ~ ( I \ s d i s t d r i b u ) t i 3 q H ~ I I C ~ I S f l a t i n t h e inid-sparl rt.": 1 on .

:n t h e o v e r 3 1 1 3t;~e~'i:ner;t of cascadt. per i o d i c i t v , qodd z i v f o i 1 - t o - j i r f o i 1 p ~ - e s s u C ? a g r e e w r l t ~ 1 s ach4euvd f3r tarit? ceriter a i ~ f o i Is s u r f a r e ~ t 3 t i f : a ~ r f c t ~ i s 4 anu 5 ) t:f th<? c?scarfes t e s t e d d t ctthsrlnic Macn n u k e r s . Tome 5 c d t t 4 ? , - ~n t o e cats .+,js o3sewsd i n t h e suct i o n - s , ~ r f a c e rocomprcssfon r e q i o n rtt t r d n s o ~ i c : and rc,p.:t-~onic e x i t "fc+ n.~,&t?rs. T t 7 i ' i w a ~ ;ff.tributeri to non- pc.- r!jdtci t y of t h e ~ , . c t iorl 5drfn.ce 5hork-bo~iri.f 3 , - y layer i n t o r i t c ! i o ~ . O a t a f o r , A ! : teczt polrlts riift 3 r d pr-esented .jnd tlsccisred i n ~ h e a ' r f o i l p r e s s l i r e d i s - t r ~ . ) , t i o n s e c t i;,ns f * ~ . the respect i v e conf i g . ~ r . d t r o n s .

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OF POOR QUALITY The d i s t r i b u t i o n o f spanwise t o t a l pressure l o s s data used t o assess mid-span two-dimensional i ty was o b t a i n e d w i t h a 19-element k i e l head rake.

These data are presented i n F i g u r e 76 f o r two subsonic Mach n u h e r s and one supersonic Mach number and show t t i a t a r e l a t i v e l y f l a t p r o f i l e was achieved between 25 percent and 75 percent of t h e span. This i n d i c a t e s t h a t reasonably good two- d i m e n s i o n a l i t y was ac4ieved. The mid-span l o s s l e v e l r z a s u r e d w i t h t h e wedge probe a t t h e same Mach n u d e r s shows good agreement w i t h t h e k i e l h e a d probe data a t t h e subsonic Mach numbers and a t t h e supersonic Mach number when t h e k i e l h e a d data a r e c o r r e c t e d f o r t h e probe bow shock.

The lack o f p e r i o d i ~ i t y i n t h e cascade f l o w f i e l d a t t r a n s o n i c and supersonic e x i t Mach ~ u r r b e r s makes p r e c i s e q u a n t i t a t i v e comparisons between data and a n a l y t i c a l p r e d i c t i o n s d i f f i c u l t . The same holds t r u e f o r t h e comparison o f data between cascades. However, t h e l e v e l o f data accuracy s t i l l p e r m i t s meaningful conclusions t o be drawn r e l a t i v e t o component development o b j e c t i v e s .

M21 PROBE

--

0 0.65 KIELHEAD 0 0 79 KIELHEAD A 1 296 KIELHEAD ADJUSTMENT FOR PROBE BOW SHOCK LOSS (M2i = 1.296)

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MID-SPAN LOSS LEVEL MEASURED WITH WEDGE PROBE

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0 1 I I 1 0 2 5 50 7 5 100 PERCENT SPAN Pressure Loss f o r Base A i r f o i l a t 0.65, F i g u r e 76 Spanwise V a r i a t i o n o f T o t a l 0.79 and 1.3 E x i t Mach Nunbers Uncooled Cascade Results -- Test and a n a l y t i c a l r e s u l t s obtained f o r t h e uncooled base, overcambered, and s t r a i g h t b a c k b l a d e designs a r e presented i n t h i s sect ion. P r e d i c t e d and measured pressure d i s t r i b u t i o n s and cascade losses are compared, and a i r f o i l t r a i l i n g edge base pressure and e x i t gas any l e a r e assessed.

4.4.3.1 Base A i r f o i 1 E v a l u a t i o n 4.4.3.1.1 - A i r f o i 1 Pressure D i s t r i b u t i o n s P r e d i c t e d and measured base a i r f o i 1 s u r f a c e s t a t i c pressure d i s t r i b u t i o n s f o r t h e range o f subsonic e x i t Mach numbers t e s t e d a r e shown i n F i g u r e s 7 7 through 79. Agreement between t h e measured subsonic data and p r e d i c t i o n s was e x c e l l e n t .

Pressure d i s t r i b u t i o n s f o r supersonic e x i t Mach numbers a r e shown i n F i g u r e s 86 through 83. Pressure d i s t r i b u t i o n s f o r o f f - d e s i g n i n l e t gas angles o f 33 and 58 degrees a t a norninal i s e n t r o p i c e x i t Mach number o f 1.3 a r e shown i n Figures 84 and 85. Agreement between p r e d i c t i o n s and t h e measured data f o r t h e supersonic e x i t c o n d i t i o n s was good except f o r t h e recompression r e g i o n where t h e measured data were s c a t t e r e d around t h e p r e d i c t i o n The v a r i a b i l i t y i n t h e data from a i r f o i l t o a i r f o i l i n t h e recompression r e g i o n I S ~ t t r i b u t e d t o n o n - p e r i o d i c i t y o f t h e f l o w f i e l d , as discussed i n Section 4.4.2 o f t h i s r e p o r t .

4.4.3.1.2 Schl i e r e n Observations S c h l i e r ;I photographs o f t h e f l o w s t r u c t u r e f o r t h e base a i r f o i l over t h e range o f i s e n t r o p i c e x i t Mach numbers t e s t e d a r e shown i n F i g u r e 86 through 89. These ~ h o t o g r a p h s i l l u s t r a t e t h a t t h e s t r e n g t h and nunher o f shocks which a f f e c t t h e l o s s increased w i t h i n c r e a s i n g Mach number. A:so, t h ~ non-period- i c i t y e f f e c t o f t h e shocks r e f l e c t e d f r o m t h e shear l a y e r can be seen.

The double r e f l e c t e d shock shown r e s u l t s f r o m t h e i n t e r a c t i o n o f a t r a i l i n g edge shock w i t h t h e s u c t i o n s u r f a c e boundary l a y e r . This i s discussed i n more d e t a i 1 i n Section 4.4.1. The double shock p a t t e r n i s most n o t i c e a b l e a t t h e higher values o f M 2 i .

T o t a l Pressure Loss Assessment The v a r i a t i o n of base b l a d e t o t a l pressure l o s s w i t h MZi a t t h e design i n l e t gas angle o f 43 degrees f o r a i r f o i l s 4 and 5 i s shown i n F i g u r e 90. The t o t a l pressure losses f o r t h e i n l e t gas angles u f 33 and 58 degrees a t M2i = 1.3 (design p o i n t Mach number) are a l s o presented i n t h i s f i g u r e and show t h a t t h e ioss changes l i t t l e w i t h incidence i v d i c a t i n g t h a t t h e design has good i n c i - dence range.

P r e d i c t e d losses which were c a l c u l a t e d employing t h e measured base pressure c o e f f i c i e n t s , CpB, a r e compared w i t h measured losses i n F i g u r e 90. This f i g u r e shows t h a t t h e p r e d i c t i o n i s i n good agreement w i t h measured losses a t subsonic Mach nurrbers and underestimate t h e losses a t supersonic Mach numbers.

TEST CONDITIONS 2 0 '63 J 2 - 1 4 5 ; M 1 - 3 2 4 6 2 1 - 53 00 l P T I'T = 5 0 1 6 3 0 AIRFOIL 4

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1 - PREDICTION

F i g u r e 77 U n c o o l o d Cascade T e s t (Base F i g u r e 78 U n c o o l e d Cascade T e s t ( B a s e B l a d e S e c t i o n ) P r e s s u r e 31 ade S e c t i o r j ) P r e s s u r e D i s t r i b u t i o n a t 0.593 E x i t D i s t r i b u t i o n a t 0.763 E x i t Mach N u h e r Mach F i u m e r 1 2 - n i0 00 1 :? 10 ) $0 9 riu 1 00 X , BX. I C A S C A D E I F i g u r e 79 I!ncooled C a s c a r l ~ Test. (Rase F i g u r e 80 U n c o o l e d Cascade T e s t (9;se 8 : :de S e c t i o n ) P r e s s u r e B l a d e S e c t i o n ) P r e s s u r e D i s t i - i b u t i o n a t 0.368 E x i t D i s t r i b u t i o n a t 1.044 E x i t .

Mach N u ~ r b e r Nach N u r k e r F i gu r e 8 1 U n c o a l e d Cascade T e s t ( B a s e F i g u r e 82 U n c o o l e d Cascade T e s t (!lase B l a d e S e c t i o n ) P r e s s u r e B l a d e S e c t i o n ) P r e s s u r e D i s t r i b u t i o n a t 1 . 1 3 7 E x i t D i s t r i b u t i o n a t 1.229 E x i t Mach N u h e r - Mach Nurrber F i g ~ ~ r e 83 U n c o o l e d Cascade T e s t f Rase F i g u r ~ 64 U n c o o l e d Cascade T e s t i Base B l a d e S e c t i o n ) P r e s s u r e B l a d e S e c t i o n ) P r e s s u r e D i s t r i b u t 'or1 a t 1.325 E x i t D i s t r i b b t ~ o n a t 1 . 2 5 1 E x i t Mach NunLl Mach N u d e r - I n l e t Gas Ang 1 e o f 33 d e g r e e s 8 1 ORIGINAL PAG? !3 OF POOR Q U A L I N F i g u r e 85 Uncooled Cascade T e s t (Base Blade S e c t i o n ) P r e s s u r e D i s t r i b u t i o n a t 1.239 E x i t Mach Number - I n l e t Gas Angle o f 58 degrees Base P r e s s u r e C o e f f i c i e n t The v a r i a t i o n o f "baseM a i r f o i l b a s e p r e s s u r e c o e f f i c i e n t , CpB, w i t h i s e n - t r o p i c e x i t Mach number, M , 1s p r e s e n t e d i n F i g u r e 91. Base p r e s s u r e c o e f - f i c i e n t s o b t a i n e d a t Mzi o f i 1 . 3 and o f f - d e s i g n i n l e t gas a n g l e s o f 33 and 58 degrees a r e a l s o p r e s e n t e d i n F i g u r e 9 1 and a r e seen t o b e a p p r o x i m a t e l y e q u a l t o t h e v a l u e o b t a i n e d a t t h e d e s i g n i n l e t gas a n g l e o f 43 decjrees. A c u r v e f a i r e d t h r o u g h t h e d a t a shows a l o c a l maximum near t h e d e s i g n e x i t Mach nurrber, w h i c h i n d i c a t e s t h a t t h e d e s i g n i n t e n t was a c h i e v e d .

4.4.3.1.5 E x i t Gas Angle - - The v a r i a t i o n o f e x i t a i r angle, P 2 , w i t h MZi f o r t h e b a s e a i r f o ; s p r e s e n t e d i n F i g u r e 92. The p r e d i c t e d e x i t a i r a n g l e i s c o n s i s t e n t l y 1 . 5 t o 2 degrees h i g h e r t h a n a c t u a l d a t a o v e r + . h e r a n g e o f Mach numbers i n v e s t i g a t e d .

The e x p l a n a t i o n f o r t h i s d i f f e r e n c e i s n o t known b u t i s r e a s o n e d t o b e r e l a t e d t o t h e measured a n g l e . C o n t i n u i t y checks and measured a i r f o i l l o a d i n g s i n d i c a - +.ed t h a t t h e p r e d i c t e d v a l u e i s more r ~ e a r l y c o r r e c t . A d j u s t i n g t h e measured a n g l e s upward b y a nominal 1.70 degrees as i n d i c a t e d by t h e s e checks, would p l a c e t h e p r e d i c t e d and a c t u a l t r e n d s w i t h Mach number i n r e a s o n a b l e agreement.

E x i t a i r a n g l e s o b t a i n e d a t t h e d e s i g n p o ' n t Mach number and o f f - d e s i g n i n l e t gas a n g l e s of 33 and 58 oegrees a r e shown i n F i g u r e 92. Changes i n i n l e t gas dnc'e would n o t b e e x p e c t e d t o have much o f an e f f e c t o n e x i t a i r anqle. T h i s p r o v e d t o b e t h e case f o r t h e i n l e t ga; a n g l e o f 58 degrees where t h e e x i t a i r a n g l e remained e s s e n t i a l l y t h e same as t h e d e s i g n e x i t a i r angle. However, F o r t h e i n l e t gas a n g l e of 33 degrees, t h e e x i t a i r a n g l e was about 2 degrezs l e s s t h a n t h e d e s i g n i n l e t a n g l e . T h i s d e v i a t i o n i s t h o u g h t t o have been caused b y e x p e r i m e n t a l i t i a c c u r a c i e s .

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-- 0 0 2 1 . .., ..- 0. I v 1 0 6 0 7 0 8 0 9 10 1 1 1 2 1 3 14 M2i F i g u r e 118 I n d i v i d u a l A i r f o i l Mass Averaged T o t a l P r e s s u r e Loss vs E x i t I s e n t r o p i c Mach N u d e r M2i F i g u r e 119 I n d i v i d u a l A i r f o i l Base P r e s s u r e vs E x i t I s e n t r o p i c Mach Number 0 6 0 7 0 8 0.9 1 0 1 1 1.2 1 3 1 4 M2i F i g u r e 120 I n d i v i d u a l A i r f o i l Mass Averaged E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Number 103 Cooled Cascade Rrsul t s Test and a n a l y t i c a l r e s u l t s o b t a i n e d f o r t h e base b l a d e w i t h t r a i l i n g edge c o o l a n t e j e c t i o n a r e presented i n t h i s section. Nominal c o o l a n t f l o w e j e c t i o n 2, and 3 percent o f cascade p r i m a r y flow, although i n one case a r a t e s were 1, f l o w r a t e o f 0.4 p e r c e n t ( F i g u r e 134) was evaluated. As was t h e case w i t h t h e uncooled cascade, p r e d i c t e d and measured pressure d i s t r i b u t i o n s and cascade losses a r e compdred and a i r f o i l t r a i l i n g edge base pressure, and e x i t gas angles assessed.

4.4.4.1 A i r f o i 1 Pressure D i s t r i b u t i o n s P r e d i c t e d and measured a i r f o i 1 s u r f a c e s t a t i c pressure d i s t r i b u t i o n s f o r t h e range o f c o o l i n g f l o w r a t e s a t subsonic e x i t Mach numbers are shown i n F i g u r e s 121 thraugh 126. Pressure distributions f o r t r a n s o n i c and supersonic e x i t Mach numbers and various c o o l i n g f l o w r a t e s are shown i n F i g u r e s 127 through 136. The p r e d i c t e d d i s t r i b u t i o n s f o r subsonic Mach numbers a r e i n e x c e l l e n t agreement w i t h t h e measured data. Agreement between t h e measured data a ~ d t h e p r e d i c t i o n s f o r supersonic Mach numbers i s g e n e r a l l y good. A comparison o f measured s t a t i c pressures obtained f o r approximately t h e same MZi and d i f f e r e n t c o o l a n t e j e c t i o n r a t e s (see F i g u r e s 122, 123, 125 and 1301 i n d i c a t e s t h a t c o o l a n t e j e c t i o n had a very small e f f e c t on t h e d i s t r i b u t i o n .

I n f a c t , t h e pressure d i s t r i b u t i o n s were n e a r l y i d e n t i c a l w i t h those obtained w i t h o u t c o o l a n t f l o w e j e c t i o n (see F i g u r e s 77 through 83).

4.4.4.2 Schl i e r e n Observations The shock s t r u c t u r e f o r t h e base a i r f o i l a t t h e design i s e n t r o p i c e x i t Mach i s presented j n F i g u r e s number and f o u r t r a i l i n g edge c o o l a n t e j e c t i o n r a t e s 136-139. A comparison o f t h e shock p a t t e r n s a t t h e d i f f e r e n t c o o l a n t e j e c t i o n r a t e s i c d i c a t e s no s i c j n i f i c a n t change i n shock p a t t e r n w i t h changes i n flow r a t e . This c o u l d l e a d t o t h e c o n c l u s i o n t h a t shock losses a r e r e l a t i v e l y un- a f f e c t e d b y t r a i l i n g edge c o o l a n t e j e c t i o n . As i s discussed i n Section 4.4.4.5, t h i s does n o t appear t o be t h e case because shock losses were found t o be n o t i c a b l y a f f e c t e d b y t r a i l i n g edge c o o l a n t e j e c t i o n , as determined from e x i t plane wake t r a v e r s e s .

4.4.4.3 Base Pressure C o e f f i c i e n t Base pressure c o e f f i c i e n t s are compared w i t h those obtained w i t h o u t c o o l a n t e j e c t i o n i n F i g u r e 140. As i n d i c a t e d , t r a i l i n g edge c o o l a n t e j e c t i o n r e s u l t e d i n base pressure c o e f f i c i e n t s which increased (became more favorable; w i t h i n c r e a s i n g t r a i i i n g edge f l o w . For example, 1 percent t r a i l i n g edge b l e e d f l o w increased t h e base pressure c o e f f i c i e n t , CpB, b y as much as 0.17 a t MZi of 1.3. I n c r e a s i n g t h e b l e e d f l o w t o 2 o r 3 percent r e s u l t e d i n approximately an a d d i t i o n a l 0.07 increase i n C p B a t M 2 i o f 1.3; r e d u c i n g CPB t o approxi- mately zero.

ORIGINAL PAGE : S OF POOR QUALI-N TEST CONDITIONS TEST C0NDITIO:dS

- 0 6 2 2 M : = 0 6 2 0

M 2

- 15 73 3 2 = 1 5 5 9

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= 0 229 M1 - 0 229

TEST CONDITIOhS M l = 0 565 = 43 00

J 1 J 1 - 4 3 0 0

= 15 8 6 1PT'F'T - 0 0 1 2 1 APTPT = O 0 1 0 9

0 9 (1 217 = 4 3 0 0 PT PT = 0 CJ99 0 3 0 7 0, a 0 a 0 AIRFOIL 4

WC W M - 0 0:667

WC W M = 0 0 2 6 2 7 0 AIRFOILS

a AIRFOIL 5 WC;WM = 0 0 1 0 3 9

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F i g u r e 1 2 1 Coo led Cascade Test Base

0 ~ 1

6 l ade Sect i on Pressure 0 3 0 2 0 0 0 0 :LO 3 4 0 0 6 0 0 9 0 1 0 0 D i s t r i b u t i o n a t 0.565 E x i t X!BX. CASCADE Mach Number TEST CONDIT'ONS TEST CONDITIONS F i g u r e 1 2 2 Cooled Cascade Test Base h l 2 = G ' 6 5 ?J 7 - ij 7 6 4 Blade S e c t i o n Pressure

= 1 5 0 8 .'2 - 15 21

j 2 D i s t r i b u t i o n a t 0.62 E x i t M1 M1 = 3 246 C 251 3 3 = 4 3 0 0 J 1 = 43 00 Mach Number 1PT,PT - 0 0 1 6 8 APTPT = 0 0 1 7 8 'EST CCXDITIONS = i) 772

0 AIRFOIL4 WC W M - 0 0 1 0 8

1 - ?WEOlCTlON

0 AIRFOIL 4 WC,WM - 0 02755

AIRFOIL 4 W C ' W M = 0 0 2 1 4

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AIRFOIL 5 W C , W U - 0 02092

0 3 1 I I I I 1 0 00 0.20 0 40 0 6 0 0 8 0 1 00 0 3 - 0 2 0 0 0 0 0 2 0 0 4 0 0 6 0 0 8 0 1.7'2 X B X . CASCADE XIRX, CASCADE ' i g u r e 123 Cooled Cascade Test Base F i g u r e 124 Cooled Cascade Test Base B lade S e c t i o n Pressure B l ade S e c t i o n Pressure D i s t r i b u t i o n a t 0.772 E x i t D i s t r i b u t i o n a t 0.765 E x i t Mach Number Mach Number ORIGINAL PAC: ;' OF POOR QIIALI I3 TEST CONDITIONS

- i 3a7

d2 - 1 4 8 0

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O A I R F O I L 4 W C I W M - 0 0 1 1 1 3 a AIRFOIL 5 WC,WM - o Ol \ I ,

4bRFOII 5 WCNiM - 0 0207

0 W 0 . 2 0 0.40 0 80 0 8 0 I 00 XiBX i i gure 126 Coo led Cascade Test Base F i g u r e 125 Coo led Cascade Test Base B l ade S e c t i o n Pressure B lade S e c t i o n Pressure D i s t r i b u t i o n a t 0.989 E x i t D i s t r i b u t i o n a t 0.98 E x i t Mach Number Mach Number TEST CONDITIONS = 1 0 2 6 = 1 4 7 8 0 9 0 = 0 258

- 4 3 00

sPT,PT = 0 0 4 9 0 0 8 0 0 7 0 I- D, om 0 5 0 0 4 0 0 AIRFOtL 4

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I PREDICTION 0 2 0 l I I I I I 0 0 0 0 2 G 0 4 0 0 6 0 0 3 0 1 0 0 X i B X F i g u r e 128 Cooled Cascade Test Base F i g u r e 127 Cooled Cascade Test Base B lade S e c t i o n P r e s s m e B lade S e c t i o n Pressure D i s t r i b u t i o n a t 1.026 E x i t D i s t r i b u t i o n at 1.034 E x i t Mach Number Mach Number TEST 30NOITIONS

- 1135

- 14 48

= 0 25; = 43.00 APT PT = 0 0 3 3 5 TEST CONOlTiONS

- 1 144

= 14 67

- t? 257

- 43 00

APT'PT = C 0332 O W 0 2 0 0.40 0 8 0 0 8 0 1 0 0 XIBX F i g u r e 130 Cooled Cascade T e s t Base F i g u r e 129 Coo l e d Cascade T e s t Base B lade S e c t i o n Pressure B l ade S e c t i o n P r e s s u r e D i s t r i b u t i o n a t 1.044 E x i t D i s t r i b u t i o n a t 1.14 E x i t Mach Number Mach Number TEST CONDITIONS

- 1 1 5 7

0 AIRFOIL 5 WC.WM 0 0 01 164

[I] AIRFOIL 5 WC/WM - 0 01 1 17

I I I 0 20 I I I I 1 0 20 I 1 0 0 0 0 2 0 0 4 0 060 0 8 0 100 0 0 0 0 2 C 0.40 0 0 0 0 8 0 100 XIBX F i g u r e 131 Coo led Cascade T e s t Base F i g u r e 132 Coo l e d Cascade T e s t Base B l ade Sect i on P r e s s u r e B lade S e c t i o n p r e s s u r e D i s t r i b u t i o n a t 1.263 E x i t D i s t r i b u t i o n a t 1,157 E x i t Mach Number Mach Number TEST CONDITIONS

- 1 273

- 1 5 8 2

- 0 258

= 4 3 0 0

APTiPT - 0 0343

0 0 0 0 2 0 0 4 0 0 6 0 0 8 0 1 0 0 XIBX F i g u r e 133 Cooled Cascade T e s t Base B l a d e S e c t i o n P r e s s u r e D i s t r i b u t i o a a t 1.273 E x i t Mach Number 1 0 TEST CONOlToONS TEST CONDITIONS

- 1 298

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,'2 = 16 38

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wc,wu - o o c u o

0 2 0 0 0 0 2 0 4 0 6 0 8 10 0 0 0 0 2 0 0 4 0 0 0 0 0 8 0 1 0 0 x:ex XIBX F i g u r e 134 Cooled Cascade T e s t Base F i g u r e 135 Coo l e d Cascade T e s t Base B lade S e c t i o n P r e s s u r e B l ade Sect i on P r e s s u r e D i s t r i b u t i o n a t 1.274 E x i t D i s t r i b u t i o s a t 1.296 E x i t Mach Number Mach Number ORIGINAL PAGE 1 9 OF POOR QUALIW F i g u r e 140 Base Pressure C o e f f i c i e n t vs Plenum E x i t I s e n t r o p i c Mach Number a t Various Cooling Flow Rates 4.4.4.4 T o t a l Pressure Loss Assessment P r e d i c t i o n s a r e compared t o measured losses, as a f u n c t i o n o f c o o l i n g f l o w t o mainstream f l o w r a t i o (Wc/Wm) and i s e n t r o p i c e x i t Mach number, i n F i g u r e s 141-146. Measured base pressure was used i n c a l c u l a t i n g t h e p r e d i c t e d values.

i n d i c a t e t h a t t o t a l pressure l o s s was i n s e n s i t i v e t o Figures 141 and 142 c o o l a n t f l o w e j e c t i o n r a t e s a t subsonic e x i t Mach numbers, t h e r e b e i n g o n l y a s l i g h t increase i n l o s s a t c o o l a n t f l o w r a t e s o f 2 and 3 percent. The p r e d i c - t i o n i n d i c a t e s t h e same t r e n d b u t overestimates t h e losses a t t h e h i g h e r f l o w r a t e s .

T o t a l pressure l o s s measurements f o r t r a n s o n i c and supersonic Mach numbers a r e shown i n Figures 143-146. I n t h e t r a n s o n i c range o f Mach numbers (1.0 t o 1.1), t h e messured losses are seen t o decrease w i t h i n c r e a s i n g c o o l a n t f l o w r a t e .

This t r e n d i s reasonably w e l l p r e d i c t e d w i t h t h e l o s s l e v e l being overestimated a t t h e higher c o o l a n t f l o w r a t e s f o r MZi = 1.0, b u t w e l l p r e d i c t e d a t MZi = 1.1. A t higher supersonic e x i t Mach nurbers, t h e measured l o s s f o r c o o l a n t f l o w r a t e s i n t h e 1-3 percent range i s e s s e n t i a l l y constant. The p r e d i c t i o n i s seen t o overestimate t h e measured losses, i n a d d i t i o n t o i n d i c a t i n g a t r e n d of i n c r e a s i n g l o s s w i t h i n c r e a s i n g c o o l a n t f l o w r a t e . The reason f o r t h i s apparent discrepancy i s discussed i n t h e f o l l o w i n g s e c t i o n .

Figure 141 Base Blade Total Pressure Loss vs Cooling Flow Ratio; M2i = 0.65 Figure 142 Base Blade Total Pressure Loss vs Cooling Flow Ratio; M2i = 0.79 -- -. -- . -- o AIRFOIL 4

T - -

0 AIRFOIL 5 I - PREDICTION i j ... - J-.- . --_J_ I I I 1 j

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- I

010 8 i ' I ! #% I i -.. .. C 0 02 I

0 I

0 0 01 0 02 0 03 WclWrn Figurr !43 Base Blade Total Pressure Loss vs Cooling Flow Ratio; Mpi = 1.00 0 10 0 AIRFOIL 4 0 AIRFOIL 5 - PREDICTION - - ~ - *--- 0.02 0 I 0 0 01 0 02 0 03 Wc: Wrn Figure 144 Base Blade Total Pressure Loss vs Cooling Flow Ratio; M z i = 1.1 0 AIRFOIL 4 0 AIRFOIL 5 0 08 Figure 145 Base Blade Total Pressure Loss vs Cooling Flow Ratio; M2i = 1.7 0 0 01 0 02 0 03 WcIWm Figure 146 Base Blade Total Pressure Loss vs Cooling Flow Ratio; M2i = 1.3 ORIGINAL B.!CL' !g OF POOR QZ;..tl'r"r 4 . 4 . 4 . 5 Shock Losses I n an i n f i n i t e cascade ( t r u l y p e r i o d i c cascade flow), t h e d i s t r i b u t i o n o f the shock and wake l o s s c o n t r i b u t i o n s t o t o t a l pressure l o s s would be as shown The experimental r e s u l t s discussed i n the pre- schematically i n Figure 147.

vious s e c t i o n i m p l i e d t h a t e i t h e r , o r b o t h o f these c o n t r i b u t o r s were a f f e c t e d b y coolant f l o w i n j e c t i o n . I n order t o v e r i f y t h i s , measured gapwise d i s t r i b u - t i o n s o f t o t a l pressure loss a t i s e n t r o p i c e x i t Mach nuntjers o f approximately 1.0 and 1.3 were examined. These measurements are shown i n Figure 148. A com- parison o f t h e d i s t r i b u t i o n s indicates t h a t shock and wake losses are indeed a f f e c t e d by t r a i l i n g edge coolant flow. A more d e t a i l e d assessment of t h e data was subsequently cond~jcted t o b e t t e r d e f i n e t h e d i s t r i b u t i o n o f shock and wake losses as impacted by changes i n coolant f l o w e j e c t i o n r a t e s and i s e n t r o p i c e x i t Mach nunher. The r e s u l t s o f t h i s assessment are shown i n Figure l d 3 , which i n d i c a t e s t h a t t r a i l i n g edge coolant f l o w e j e c t i o n causes an increase i n t h e decrease i n shock loss dominating wake loss and a decrease i n shock loss; and causing a decrease i n t o t a l pressure loss. This observation i s confirmed It would appear then, t h a t t h e coolant f l o w b y a r e - e x m i n a t i o n o f Figure 148.

i n j e c t e d i n t o t h e t r a i l ~ n g edge f l o w f i e l d serves t o weaken t h e t r a i l i n g edge shocks, thereby reducing the losses associated w i t h t h e shock s t r u c t u r e and, q u i t e possibly, those associated w i t h shock-boundary l a y e r i n t e r a c t i o n s , although the l a t t e r cannot be d i r e c t l y shown b y t h e data presented.

As mentioned e a r l i e r , the Schlieren v i s u a l i z a t i o n s d i d not i n d i c a t e a percept- able change i n t r a i l i n g edge shock p a t t e r n when coolant f l o w was injected.

This seems t o be i n d i r e c t c o n t r a s t w i t h the l o s s r e s u l t s j u s t discussed.

However, i t i s hypothesized t h a t t h e shocks shown i n t h e Schlieren v i s u a l i z a - t i o n s formed a t the r i b s between t h e t r a i l i n g edge coolant e j e c t i o n holes as shown i n F i ure 150. This would r e s u l t i n an a l t e r n a t i n g zone o f shocks and shock-free 9 o r weak shock) layers along t h e t r a i l i n g edge of the blade span.

To the Schlieren system, these shock waves would appear i d e n t i c a l t o those eminating from a s o l i d t r a i l i n g edge. The presence o f these "layered" shocks may have caused some o f t h e "between-wake" t o t a l pressure loss i n d i c a t e d a t MZi = 1.3 i n Figure 148. A t M = 1.0, these shocks would be too weak t o have a n o t i c a b l e e f f e c t on t o Pi a1 pressure loss.

I 4 AIRFOIL NO.

Figure 147 Schematic o f Shock and Wake Loss C o n t r i b u t i o n s t o Total Pressure Loss WclWm = .0212 M2i = 1.302 WclWm = 0.0 M2i = 0.0994 F i g u r e 148 Gapwise D i s t r i b u t i o n s o f T o t a l P r e s s u r e Loss For Base A i r f o i l s With S e v e r a l C o o l i n g Flow Rates And I s e n t r o p i c Mach Numbers OF 1.0 And 1.3 ORIGINAL PAGE IS OF POOR QUALITY F i g u r e 149 T o t a l Pressure, Shock And Wake Loss vs E x i t I s e n t r o p i c Mach Nunher a t Various C o o l i n g Flow Rates F i g u r e 150 Schematic of Shock P a t t e r n Formed a t Ribs Between T r a i l i n g Edge Coolant E j e c t i o n Holes 4.4.4.6 E x i t Gas Angle The v a r i a t i o n o f e x i t a i r angle w i t h M e i a t several c o o l a n t flow e j e c t i o n r a t e s i s compared w i t h data obtained w i t h no f l o w e j e c t i o n and p r e d i c t e d a i r angles i n F i g u r e 151. Generally, t h e e x i t angles o b t a i n e d f o r f l o w e j e c t i o n r a t e s o f 1.0, 2.0 and 3.0 percent were i d e n t i c a l and s i m i l a r t o those o b t a i n e d w i t h o u t f l o w e j e c t i o n .

P r e d i c t i o n s overestimate measured subsonic and t r a n s o n i c data b y approximately 1.5 degrees. The e x p l a n a t i o n f o r t h i s d i f f e r e n c e i s t h e satre as t h a t f o r t h e discrepancies noted i n t h e base and overcanbered a i r f o i l e x i t gas angle data discussed i n Sections 4.4.3.1.5 and 4.4.3.2.5 o f t h i s r e p o r t . I f t h i s apparent discrepancy i s removed, t h e r e i s good agreemerib between measured and p r e d i c t e d values up t o MZi = 1.1. Beyond t h a t , agreement becomes poor.

4.4.5 Summary o f Blade Cascade Results - Results o f b l a d e cascade t e s t i n g are sunlmarized i n t h e f o l l o w i n g sections.

This summary compares t h e measured t o t a l pressure loss, base pressure c o e f f i - c i e n t , and e x i t a i r angles f o r t h e base, overcanbered, and s t r a i g h t b a c k a i r - f o i l s . It a l s o presevts 3 comparison o f t h e base a i r f o i 1 performance w i t h and w i t h o u t t r a i l i n g edge c o o l a n t f l o w e j e c t i o n .

4.4.5.1 Comparison of Uncooled Base, Overcambered, and Straightback A i r f o i l s 4.4.5.1.1 T o t a l Pressure Loss T o t a l pressure losses, A P T / P ~ , f o r t h e base, overcanbered, and s t r a i g h t .

back a i r f o i l s are summa;-ized i n F i g u r e 152 as a f u n c t i o n of M2i. Also i n - cluded are base a i r f o i l pressure losses a t design p o i n t M 2 i f o r off-design i n l e t gas angles o f 33 and 58 degrees. The curves shown a r e f a i r e d through averaged data.

These data i n d i c a t e r e s u l t s c o n s i s t e n t w i t h t h e base pressure c o e f f i c i e n t r e s u l t s , i .e., t h e base a i r f o i 1 provides t h e lowest t o t a l pressure losses f o r subsonic and low supersonic M and t h e pressllre losses f o r a l l e t h r e e a i r f o i l s are e s s e n t i a l l y equal a t he design p o i n t M2i. Base a i r f o i l pressure l o s s appears t o b e r e l a t i v e l y i n s e n s i t i v e t o cl-,anges i n i n l e t gas angle over t h e range shown, i n d i c a t i n g t h a t t h e base b l a d e design has good incidence range capabi l i t y .

4.4.5.1.2 Base Pressure C o e f f i c i e n t The v a r i a t i o n o f base, overcambered and s t r a i g h t b a c k a i r f o i l base pressurz c o e f f i c i e n t s , CpB, w i t h i s e n t r o p i c e x i t Mach number (MZi) i s summarized i n F i g u r e 153. Curves have been f a i r e d through t h e averaged d a t a measured from each cascade. As i n d i c a t e d , t h e base a i r f o i l p r o v i d e s t h e most f a v o r a b l e o v e r a l l v a r i a t i o n o f CPB w i t h M2i, although a t t h e design p o i n t e x i t Mach number t h e r e s u l t s f o r a l l t h r e e a i r f o i l s a r e close. As noted e a r l i e r , t h e s t r a i g h t b a c k a i r f o i l design appears t o o f f e r a more f a v o r a b l e CP than t h e

I ! base o r overcamber designs a t MZi g r e a t e r than 1.3, b u t t h i s Mac number

r e g i o n i s g r e a t e r than t h e component design p o i n t Mach number.

ORIGINAL OF POOR FROM FIG 921 OPEN SYMBOLS - AIRFOIL 5

I

: , I \'ill PlJ'VT SLASHED SYMBOLS - AIRFCIL 4 0 I F i g u r e 151 Base B l a d e E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Number a t V a r i o u s C o o l i n g Flow Rate - o BASE I I --- O OVERCAMBER -- A STRAIGHT BACK 0 BASE (INLET GAS ANGLE = 33 DEGREES F i g u r e 1 5 2 Comparison o f T o t a l P r e s s u r e Loss vs E x i ~ I s e r l t r c p i c Mach Nunher F i g u r e 153 Comparison of Base P r e s s u r e vs E x i t I s e n t r o p i c Mach Number 4.4.5.1.3 - E x i t A i r Angle The v a r i a t i o n o f e x i t a i r angle, p 2 , w i t h MZi f o r t h e base, overcambered, and s t r a i g h t b a c k a i r f o i l s i s compared w i t h p r e d i c t e d P2 i n F i g u r e 154. I n general, curves f a i r e d through t h e naveraged46 data show t h e same t r e n d , d i f - f e r i n g o n l y i n l e v e l . If t h e base and overcamber a i r f o i l data are a d j u s t e d up- ward b y approximately 1.7 and 1.2 degrees r e s p e c t i v e l y , based on c o n t i n u i t y checks and a i r f o i l loading c a l c u l a t i o n s , t h e e x i t a i r angles f o r a l l t h r e e designs become approximately equal and i n reasonable agreement w i t h t h e p r e d i c t i o n .

4.4.5.2 C o m ~ a r i s o n Between Cooled and Uncooled Base A i r f o i 1s 4.4.5.2.1 T o t a l Pressure Loss T o t a l pressure losses obtained f o r t h e base a i r f o i l w i t h t r a i l i n g edge c o o l a n t f l o w e j e c t i o n r a t e s o f approximately 1, 2 and 3 p e r c e n t a r e compared w i t h l o s s i n F i g u r e 155. In general t r a i l i n g edge f l o w o b t a i n e d w i t h no f l o w e j e c t i o n e j e c t i o n causes a small increase i n t o t a l pressure l o s s a t subsonic Mach numbers b u t reduces t o t a l pressure l o s s a t supersonic Mach numbers. The data i n d i c a t e t h a t f l o w e j e c t i o n i s most e f f e c t i v e i n r e d u c i n g t o t a l pressure l o s s f o r M2i g r e a t e r than 1.1. This i s a t t r i b u t a b l e t o t h e f a c t t h a t f l o w e j e c - t i o n weakens t h e t r a i l i n g edge shock s t r e n g t h such t h a t decreases i n shock losses predominate over corresponding increases i n wake l o s s caused b y t h e added f l o w a t t h e t r a i l i n g edge.

Base Pressure C o e f f i c i e n t T r a i 1 i n g edge c o o l a n t f l o w e j e c t i o n r e s u l t e d i n base pressure c o e f f i c i e n t s which increased (became more f a v o r a b l e ) w i t h i n c r e a s i n g t r a i 1 i n g edge f l o w and r e f l e c t e d t h e decrease i n t o t a l pressure l o s s !see F i g u r e 140). The data i n d i c a t e t h a t 1 percent t r a i l i n g edge f l o w e j e c t i o n increased t h e base pressure c o e f f i c i e n t , CpB, by as much as 0.17 a t M2i o f 1.3. I n c r e a s i n g t h e f l o w t o 2 o r 3 percent r e s u l t e d i n approximately an a d d i t i o n a l 0.07 increase i n C p B a t Mzi of 1.3; reducing C p B t o approximately zero.

L i t A i r Angle The v a r i a t i o n i n e x i t a i r angle, p 2 , w i t h v a r i a t i o n s i n MZi a t s e v e r a l t r a i l i n g edge c o o l a n t f l o w e j e c t i o n r a t e s i s compared t o p r e d i c t i o n s and d a t z o b t a i n e d w i t h no flow e j e c t i o n i n F i g u r e 156.

The d a t a show t h a t t r a i l i n g edge f l o w e j e c t i o n had l i t t l e e f f e c t on e x i t a i r angle. P r e d i c t i o n s agreed w i t h t h e data trends ~ u t tended t o o v e r - p r e d i c t P 2 b y t1.0 t o -2.0 degrees. This discrepancy i s unexplained as discussed i n prev?ous s e c t i o n s of t h i s r e p o r t .

C o n t i n u i t y and a i r f o i 1 loading c a l c u l a t i o n s i n d i c a t e d t h a t t h e data should be a d j u s t e d upward b y approximately 1.5 degrees t o c o r r e c t t h e e r r o r . T h i s adjustment b r i n g s t h e data and p r e d i c t i o n i n t o reasonable agreement w i t h each o t h e r .

P A G X S ORIGINAL QUALITY OF POOR F i g u r e 154 Comparison o f E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Number F i g u r e 155 Comparison o f T o t a l P r e s s u r e Loss vs E x i t I s e n t r o p i c Mach Number F i g u r e 156 Compar:sdn of E x i t A i r Angle vs E x i t I s e n t r o p i c Mach Number (Average of A i r f o i I s 4 and 5 ) 5.0 CONCLUSIONS 5.1 Vane Cascade o The performance b e n e f i t s o f t h e contoured vane endwall d e s i ~ o have been confirmed f o r t h e high-pressure t u r b i n e component. The S-wall cascade has 1 7 percent l e s s full-passage, mass-averaged pressure l o s s than t h e s t r a i g h t w a l l cascade.

o I n t h e two-dimensional spanwise f l o w r e g i o n o f t h e cascade, vane e x i t a i r angle i s g e n e r a l l y i n s e n s i t i v e t o changes i n endwall c o n f i g u - r a t i o n , v a r i a t i o n s i n e x i t Mach number, and v a r i a t i o n i n c o o l a n t f l o w r a t e s .

o V a r i a t i o n s i n vane e x i t Mach number, i n b o t h t h e c ~ o l e d and uncooled s t r a i g h t w a l l cascades, i n d i c a t e d no abrupt pressure l o s s i n c r e a s e ( t r a n s o n i c drag r i s e ) even a t near-sonic c o n d i t i o ~ ~ s . This confirms t h a t t h e base vane s e c t i o n i s acceptable f o r t h e component des2gn.

o Mid-span t o t a l pressure l o s s almost doubled when d e s i g n - p o i n t c o o l a n t f i l m c o o l i n g s i t e s . The l a r g e s t c o n t r i b u t o r f l o w was e j e c t e d f r o a a l l t o t h i s increased pressure l o s s was s u c t i o n s u r f a c e c o o l a n t f l o w i n j e c t i o n , showing a p e n a l t y abaut 5 times h i g h e r t h a n t h a t f o r pressure s u r f a c e i n j e c t i o n . T r a i 1 i n g edge design poi!;: c o o l a n t f l o w i n j e c t i o n had 1 i t t l e e f f e c t on cascade performance.

o P r e d i c t e d a i r f o i l pressure d i s t r i b u t i o n s were i n good jgreement w i t h measured s t a t i c pressures f o r b o t h two- and three-dimensional cascade f l o w c o n d i t i o n s over t h e r?nge o f e x i t Math n u h e r s and c u o l a n t f l o w c o n d i t i o n s evaluated.

5.2 Blade Cascades o Testing v e r i f i e d t h e d i s t r i b u t i o n o f c u r v a t u r e s e l e c t e d f o r t h e high-pressure t u r b .se component b l a d e a i r f o i 1 geometry. 'This geometry (base a i r f o i 1 ) provided t h e lowest t o t a l .pressure losses f o r subsonic and t r a n s o n i c e x i t Mach numbers, and pressure losses, f r o q an e x p e r l - mental s t a n d p o i n t , equal t o t h e o t h e r , a i r f o i l s a t t h e design p o i n t e x i t Mach number.

o Base a i r f o i l pressure l o s s was r e l a t i v e l y i n s e n s i t i v e t o v a r i a t i o n s i n i n l e t gas angle over a range o f 2 5 degrees, i n d i c a t i n g t h a t t h e component b l a d e design nas good incidence range.

o Testing o f t h e base b l a d e cascade w i t h t r a i l i n g edge c o o l a n t f l o w e j e c t i o n i n d i c a t e d t h a t t h i s caused a small i n c r e a s e i n pressure l o s s a t subsonic Mach numbers, b u t reduced t o t a l pressure l o s s a t super- sonic e x i t Mach numbers r e l a t i v e t o t h e base b l a d e w i t h o u t c o o l a n t f l o w e j e c t i o n . This behavior i s a t t r i b u t a b l e t o t h e f a c t t h a t f l o w e j e c t i c n weakens t h e t r a i 1 i n g edge shock s t r e n g t h such t h a t decreases i n shock iosses predominate over corresponding increases i n wake l o s s caused b y t h e added f l o w a t t h e t r a i l i n g edge.

o Predicted a i r f o i l pressure d i s t r i b u t i o n s were i n good aqreement w i t h measured d i s t r i b u t i o n s f o r b o t h cooled and uncooled cascades a t sub- sonic e x i t Mach nuthers. Agreement between the predicted and measured data f o r supersonic e x i t conditions was good except f o r the s u c t i o n surface recompression region, where the data were scattered arcund the prediction. This data s c a t t e r i s a t t r i b u t e d t o t h e nonperiodic nature o f the suction surface shock-boundary l a y e r i n t e r a c t i o n i n the cascades.

o For the th. ee uncooled cascades, e x i t a i r angle vs Mach number trends were s i m i l s r and reasonably w e l l predicted. However, experimental d i f f i c u l t i e s l e d t o unexplained differences between p r e d i c t i o n s and measured e x i t a i r angle data f o r t h e base and overcambered uncooled cascades and the base cascade w i t h t r a i l i n g edge coolant f l o w e j e c - t i o n . Correcting f o r these d i f f e r e n c e s through t h e use o f c o n t i n u i t y and a i r f o i 1 loading c a l c u l a t i o n s i n d i c a t e d t h a t the p r e d i c t i o n s were more n e a r l y c o r r e c t and t h a t t h e three uncooled c o n f i g u r a t i o n s achieve approximately the same e x i t a i r angles. The p r e d i c t e d and measured data f o r a l l of t h e cascades were 'n reasonably good agree- ment. The a d d i t i o n o f t r a i l i n g edge coolant f l o w e j e c t i o n !tad l i t t i e e f f o c t on e x i t a i r angle.

APPENDIX A The f o l l o w i n g appendices 1 i s t t h e coordinatss c f t h e vane and b l a d e a i r f o i l s employed i n t h e cascade t e s t i n g described i n t h i s r e p o r t . F i g u r e s A - 1 and A-2 d e f i n e t h e a p p l i c a b l e nomenclature.

ORIGINAL PP.32 i3 I t e l c,\- ; - - * I f

OF POOR C

Y X Figure A-1 Nomenclature Used to Define Airfoil Coordinates Figure A-2 Nomenclature Used to Define Profiled Wall Coordinates in S-Wall Vane Cascade TABLE A - 1 VANE AIRFOIL COORDINATES FOR S-WALL 3ASCADE Cascade a x i a l c h o r d , Bx = 3.34 cm (1.315 i n ) TABLE A-2 VANE A i R F O I L COORDINATES FOR STRAIGHT WALL CASCADE Cascade a x i a l c h o r d , 6, = 3.46 crn (1.364 i n ) TABLE A - 3 BASE BLADE A I R F O I L COORDINATES Cascade axial chord, Bx = 2.34 cm ( 0 . 9 2 2 i n ) TABLE A-4 OVERCAMBERED BLADE A I R F O I L COORDINATES Cascade a r i a l chord, Bx = 2 . 3 4 cm (0.922 i n ) TABLE A-5 STRAIGHTBACK BLADE AIRFOIL COORDINATES Cascade a x i a l chord, Bx = 2.34 cm (0.922 i n )

APPENDIX B LIST O F SYMROLS Annulus area A R X A x i a l chord

Base pressure c o e f f i c i e n t , Cps = (PB - P2) /Q2

C P O A x i a l f l o w v e l o c i t y c x I n c i d e n c e angle i I s e n t r o p i c e x i t Mach number M2 i E x i t Mach number M2 N Mechanical speed, rev./mi n Base s t a t i c pressure ( a t t r a i l i n g edge) P B S t a t i c pressure

p s

T o t a l pressure

" T

P Free stream s t a t i c pressure Upstream t o t a l pressure PTO Downstream f r e e stream dynamic pressure, Q = 1/2 P V?

Reynolds number based on a x i a l chord and e x i t c o n d i t i o n s

re^^

Cool ant t o t a l temperature Toc Primary f l o w t o t a l temperature To P U Tangent i a1 wheel speed Free stream v e l o c i t y root-mean-square value o f v e l o c i t y CL' f l u c t u a t i o n i n t h e streamwise d i r e c t i n n Coolant mass flow Cascade p r i m a r y (mainstream) mass f l o w A x i a l d i r e c t i o n i n t u r b i n e C i r c u m f e r e n t i a l d i r e c t i o n i n t u r b i n e L I S T OF SYMBOLS (Cont ' d l (Y Vane a i r angle measured from t a n g e n t i a1 suhscr i p t s

P Blade a i r angle measured f r o m t a n q e n t i a l s u o s c r i p t s

A Incremental value

r Uncovered t u r n i n g

s u b s c r i p t s 1 Upstream r e f e r e n c e plane 2 Downstream measurement plane Downstream m i xed o u t t o homogeneous c o n d i t i o n s plane REFERENCES 1. Deich, M. E., Zaryankin, A. E., F i l l i p o v , G . A , , and Zatsepin, M. F., "Method o f I n c r e a s i n g t h e E f f i c i e n c y o f Turbine Stages w i t h Short Blades," Teploenergetika, Vol. 2, February, 1960. pp. 240-254.

2. Ewen, J. S . , Huber, F. W., and M i t c h e l l , J. P., " I n v e s t i g a t i o n o f t h e Aerodynamic Performance o f Small Axi a1 Turbines, " ASME Paper No. 73-GT-3, Washington, D. C., A p r i l 1973.

3. M o r r i s , A. W. H. and Hoare, R. G., ''Secondary Loss Measurements i n a Cascade o f Turbine Blades w i t h M e r i d i o n a l Wall P r o f i 1 ing," ASME Paper No.

75-WA/GT-13, Houston, Texas, Noverrber 1975.

4. Came, P. M., "Secondary Loss Measurements i n a Cascade o f Turbine Blades," I n s t i t u t e o f Mechanical Engineers Conference P u b l i c a t i o n s No. 3, Heat and F l u i d F l o w i n Steam and Gas Turbine Plant, Warwick, 1973.