with mach number inlets - ntrs.nasa.gov · high mach number inlets edwmd lunzsdaine, jenn g....

N A S A

00 0 h cv x U

I

C O N T R A

R E P O R T C T O.R=

LOAN COPY: RETURN TO AFWL TECPlNBCAL LIBRARY

KIRTLAND AFB, N. M.

NOISE SUPPRESSION WITH HIGH MACH NUMBER INLETS

Edwmd Lunzsdaine, Jenn G. Cherng, and Ismail Tag

Prepared by T-L- TENNESSEE u, : I y . 1 ,

ij I, i. . I a, r Knoxville, Tenn. 37916 1 .,. , , , , '; , :, : , , ., 1

f o r Langley Research Center :"/ , 1- I . - ?

I , . ! ' , . ' g>.,: , ' I

N A T I Q N A L A E R O N A U T I C S A N D S P A C E A D M I N I S T R A T I O N W

https://ntrs.nasa.gov/search.jsp?R=19760021871 2019-02-03T22:17:19+00:00Z

TECH LIBRARY KAFB. NM

" " ~~ ~

1. Report No. I 2. Government Accerrion No. - ..

NASA CR-2708

4. Title and Subtitle -:. ~~ I

NOISE SUPPRESSION WITH HIGH MACH NUMBER INLETS

7. Author(s) . -~ "" .

Edward Lumsdaine, Jenn G. Cherng, and Ismail Tag

-~ ~ _ _ _ ~ ~ "" ~-~ 9. Performing Organization Name and Address

The U n i v e r s i t y o f Tennessee Department o f Mechan ica l and Aerospace Engineering Knoxvi 1 l e . Tennessee 37916

~~

12. Sponsoring Agency Name and Address

Nat ional Aeronaut ics and Space A d m i n i s t r a t i o n Washington, D. C. 20546

1

3. Recipient's C a t a l o g No.

5. Report Date August 1976

6. Performing Organization Code

8. Performing Organization Report No.

10. Work Unit No.

11. Contract or Grant No.

13. Type of Report and Period Covered

C o n t r a c t o r r e p o r t

14. Sponsoring Agency Code

I

15. Supplementary Notes The au tho rs w i d l i k e t o acknowledge the ass is tance g iven to th is p rogram by Mr. .orenzo R . C la rk o f t he Acous t i cs B ranch , Acous t i cs and Noise Reduct ion Div is ion, Langley Research :en ter . M r . C la rk was r e s p o n s i b l e f o r c o o r d i n a t i n g t h e o v e r a l l e x p e r i m e n t a l e f f o r t . T h i s i n c l u d e d s e t ~p o f t e s t hardware and inst rumentat ion, operat ion o f the Langley no ise research compressor , and : o l l e c t i o n and r e d u c t i o n o f d a t a . These a c t i v i t i e s were e s s e n t i a l t o t h e c o m p l e t i o n o f t h e r e s e a r c h

.6 . Abst ract jrogram and-tre. great!Y-aP=t%dL - . ~ . _ ._

Exper imenta l resu l ts have been ob ta ined fo r two types o f h igh Mach number i n l e t s , one w i t h a t r a n s l a t i n g c e n t e r b o d y and a f i x e d g e o m e t r y i n l e t ( c o l l a p s i n g c o w l ) w i t h n o c e n t e r b o d y . A s tudy was made o f t h e a e r o d y n a m i c and a c o u s t i c p e r f o r m a n c e o f t h e s e i n l e t s . The e f f e c t s o f a r e a r a t i o , l e n g t h / d i a m e t e r r a t i o , and l i p geometry were among severa l parameters inves t iga ted . The t r a n s l a t i n g c e n t e r b o d y t y p e i n l e t was found t o be s u p e r i o r t o t h e c o l l a p s i n g c o w l b o t h a c o u s t i c a l l y a n d a e r o d y n a m i c a l l y , p a r t i c u l a r l y f o r a r e a r a t i o s g r e a t e r t h a n 1.5. Comparison o f l e n g t h / d i a m e t e r r a t i o and a r e a r a t i o e f f e c t s on performance near choked f low showed the l a t t e r t o be more s i g n i f i c a n t . A l s o , g r e a t e r h i g h f r e q u e n c y n o i s e a t t e n u a t i o n was achieved b y i n c r e a s i n g Mach number f rom low to h igh subson ic va lues .

17. Key Words (Suggested by Author(s)) . ~ .

18. Distribution Statement

A c o u s t i c s , i n l e t , c o m p r e s s o r , n o i s e a t t e n t u a t i o n , aerodynamic per formance, choked f low Unclass i f ied - U n l i m i t e d

Subject Category 7 1 " ~ - ~~ "~ ~ ~"

19. Security Classif. (of this report1 22. Price' 21. No. of Pages 20. Security Classif. (of this page)

U n c l a s s i f i e d $5.25 1-06 U n c l a s s i f i e d "

* For sale by the National Technical Information Service, Springfield, Virginia 22161

CONTENTS

SUMMARY

LIST OF ILLUSTRATIONS

1. INTRODUCTION

2 . DESCRIPTION OF INLET CONFIGURATIONS

Table I

3. PROCEDURE FOR INLET DESIGN

4 . TEST FACILITY, INSTRUMENTATION AND PROCEDURE

Page

iv

V

1

4

5

6

8

5. DISCUSSION OF RESULTS 11

6. COMPARISON OF RESULTS 1 4

7. EMPIRICAL EQUATIONS TO ESTIMATE NOISE ATTENUATION 17

8. CONCLUSION 20

9 . REFERENCES 22

FIGURES 24

APPENDIX: REVIEW OF WORK ON SONIC AND NEAR-SONIC INLETS (TABLE A) 99

iii

SUMMARY

A study was made of the parameters affecting the aerodynamic and acoustic

performance of high Mach number inlets using the translating centerbody and

fixed geometry configurations. The study included the effects of area ratio,

length/diameter ratio, and lip geometry on the acoustic and aerodynamic performance when the rotor is at subsonic as well as supersonic tip speed.

The results support earlier findtngs by the authors that the translating

centerbody type inlet is superior to the collapsing cowl, both acousti-

cally and aerodynamically, especially at moderately high area ratios

(Aexit’Athroat > 1.5). The length/diameter ratio does not seem to be as crucial to performance near choked flow as area ratio. Inlets operacing

at high Mach numbers are more effective in reducing high frequency noise.

At choked flow, however, the low frequency noise is also effectively reduced.

This study also showed that the actual amount of noise reduction depends on the flow downstream of the throat (pressure recovery) in contradiction to

inviscid theory. Choking does not guarantee a large amount of noise reduction

if it is accompanied by high pressure loss. Thus, without boundary layer

control, choked inlets are area ratio limited.

Using the present test results, an empirical formula was derived, relating

the noise reduction to percent of maximum mass flow and pressure recovery.

Results of previous studies, where noise attenuation was related to throat

Mach number, are ambiguous and are difficult to assimilate for comparative

purposes. The percent of maximum mass flow is a more satisfactory parameter,

although accuracy in measurement is most vital.

Simulating forward speed by contouring the lips of the inlets was difficult,

and the results were inconclusive. However, the lip design had a significant

effect on the aerodynamic and acoustic performance of the inlets.

iv

LIST OF ILLUSTRATIONS

Figure 1

2

3

4

5

6

7

8

9

10

11

12

13

14

WALL AND CENTERLINE MACH NUMBERS FOR POTENTIAL E'LOW IN TWO DIMENSIONS

INFLUENCE OF LENGTH/DIAMF,TER RATIO ON NOISE ATTENUATION (EJECTOR SOURCE)

SCHEMATIC AND AREA DISTRIBUTION OF TRANSLATING CENTER- BODY INLET CONFIGURATION 1


SCHEMATIC AND AREA DISTRIBUTION OF TRANSLATING CENTER- BODY INLET CONFIGUKATION 3


SCHEMATIC AND AREA DISTRIBUTION OF FIXED GEOMETRY (COLLAPSING COWL) INLET CONFIGURATION 5

SCHEMATIC AND AREA DISTRIBUTION OF FIXED GEOMETRY (COLLAPSING COWL) INLET CONFIGURATION 6

SCHEMATIC AND AREA DISTRIBUTION OF FIXED GEOMETRY INLET CONFIGURATIONS 7,8,9

AERODYNAMIC AND ACOUSTIC PERFORMANCE OF INLET WITH DIFFERENT LENGTH/DIAMETER RATIOS

OPTIMUM L/D FOR A GIVEN AREA RATIO AT DIFFERENT THROAT MACH NUMBERS FOR CIRCULAR AND SQUARE DIFFUSERS

OPTT" L/D FOR A GIVEN AREA RATIO AT DIFFERENT THROAT MACH NUMBERS FOR ANNULAR DIFFUSERS

THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 0 CM DISPLACEMENT

THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS

Page

2 4

25

26

27

28

29

30

31

32

33

34

35

36

FOR CONFIGURATION 1 W I T H CENTERBODY AT 5.1 CM DISPLACEMENT 37

V

Figure

1 5 THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 12.7 CM DISPLACEMENT

16 THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 1 7 . 8 CM DISPLACEMENT

17

1 8

1 9

20

2 1

22

23

24

25

26

27

28

29

30

3 1

32

33

34

THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 20.3 CM DISPLACEMENT

NASA-LANGLEY RESEARCH CENTER ANECHOIC-CHAMBER TRANSONIC COMPRESSOR FACILITY

INSTRUMENTATION ROOM, ANECHOIC-CHAMBER TRANSONIC COMPRESSOR FACILITY, NASA-LANGLEY RESEARCH CENTER

SCHEMATIC OF INLET CONFIGURATIONS 2 , 3 AND 4 WITH AERODYNAMIC INSTRUMENTATION

SCHEMATIC OF INLET CONFIGURATION 5 WITH AERODYNAMIC INSTRUMENTATION

ILLUSTRATION OF INLET AND COMPRESSOR OPERATINC CONDITIONS FOR DATA ACQUISITION

NOISE ATTENUATION FOR INLET CONFIGURATION 1

MASS FLOW FOR INLET CONFIGURATION 1

AVERAGE MAXIMUM MACH NUMBER OF INLET CONFIGURATION 1

PRESSURE RECOVERY FOR INLET CONFIGURATION 1

PRESSURE DISTORTION FOR INLET CONFIGURATION 1



AVERAGE MAXIMUM MACH NUMBER OF INLET CONFIGURATION 2

PRESSURE RECOVERY FOR INLET CONFIGURATION 2

PRESSURE DISTORTION FOR INLET CONFIGURATION 2



38

39

40

4 1

42

43

4 4

45

46

47

48

49

50

5 1

52

53

54

55

56

57

vi

Figure 35

36

37

38

39

40

4 1

42

43

44

45

46

47

48

49

50

51

52

53

54

55

56

57

Page

AVERAGE MAXIMUM MACH NUMBER FOR INLET CONFIGURATION 3 58

PRESSURE RECOVERY FOR INLET CONFIGURATION 3 59

PRESSURE DISTORTION FOR INLET CONFIGURATION 3 60

NOISE ATTENUATION FOR INLET CONFIGURATION 4 6 1

MASS FLOW FOR INLET CONFIGURATION 4 62

AVERAGE MAXIMUM MACH NUMBER FOR INLET CONFIGURATION 4 63

PRESSURE RECOVERY FOR INLET CONFIGURATION 4 64

PRESSURE DISTORTION FOR INLET CONFIGURATION 4 65

EFFECT OF PERCENTAGE OF MAXIMUM MASS now ON NOISE ATTENUATION FOR INLET CONFIGURATION 2 66

EFFECT OF PERCENTAGE OF MAXI" MASS FLOW ON NOISE ATTENUATION FOR INLET CONFIGURATION 3 67

EFFECT OF PERCENTAGE OF MAXIMUM MASS FLOW ON NOISE ATTENUATION FOR INLET CONFIGURATION 4 68

NOISE ATTENUATION FOR INLET CONFIGURATIONS 5AND 6 69

MASS FLOW FOR INLET CONFIGURATIONS 5AND 6 70

PEAK WALL MACH NUMBER OF INLET CONFIGURATIONS 5AND 6 7 1

PRESSURE RECOVERY FOR INLET CONFIGURATIONS 5AND 6 72

PRESSURE DISTORTION FOR INLET CONFIGURATIONS 5 AND 6 7 3

EFFECT OF PERCENTAGE OF MAXIMUM MASS FLOW ON NOISE ATTENUATION FOR INLET CONFIGURATIONS 5 AND 6 7 4

MASS FLOW FOR INLET CONFIGURATIONS 7, 8 AND 9 75

PEAK WALL MACH NUMBER OF INLET CONFIGURATIONS 7, 8 AND 9 76

PRESSURE RECOVERY FOR INLET CONFIGURATIONS 7, 8 AND 9 77

PRESSURE DISTORTION FOR CONFIGURATIONS 7 , 8 , 9 7 8

COMPARISON OF OVERALL NOISE ATTENUATION CHARACTERISTICS OF INLETS OPERATING AT SUBSONIC MACH NUMBERS 79

COMPARISON OF BLADE PASSAGE FREQUENCY ATTENUATION CHARACTERISTICS OF INLETS OPERATING AT SUBSONIC MACH NUMBERS 80

v i i

Figure

58

Page

81 FREQUENCY SPECTRA FOR CONFIGURATION 2 AT 17,500 RPM

59 82 FREQUENCY SPECTRA FOR CONFIGURATION 2 AT 20,000 RPM

60

6 1

INFLUENCE OF L I P GEOMETRY ON NOISE ATTENUATION 83

INFLUENCE OF L I P GEOMETRY ON NOISE ATTENUATION AND PRESSURE RECOVERY 84

6 2 INFLUENCE OF L I P GEOMETRY ON NOISE ATTENUATION AND PRESSURE RECOVERY 85

INFLUENCE OF L I P GEOMETRY ON NOISE ATTENUATION AND PRESSURE RECOVERY

63 86

64 INFLUENCE OF L I P GEOMETRY ON NOISE ATTENUATION AND PRESSURF, RECOVERY 87

88 65

66

INFLUENCE OF CENTERBODY ON NOISE LEVEL

POLAR GRAPHS SHOWING THE INFLUENCE OF L I P GEOMETRY OF INLETS OPERATING AT SUBSONIC MACH NUMBER 89

90 67

68

INFLUENCE OF COMPRESSOR PRESSURE RATIO ON NOISE

COMPARISON OF NOISE ATTENUATION CHARACTERISTICS OF INLETS OPERATING AT HIGH SUBSONIC MACH NUMBER WITH COMPRESSOR OPERATING AT SUPERSONIC ROTOR T I P SPEED 9 1

69 EFFECT OF AREA RATIO ON NOISE ATTENUATION AND PRESSURE RECOVERY 9 2

7 0 INFLUENCE OF PRESSURE RECOVERY ON THE EXPONENTIAL PARAMETER a 93

7 1

7 2

73

7 4

COMPARISON BETWEEN EMPIRICAL EQUATIONS AND EJECTOR TEST DATA 94

COMPARISON BETWEEN EMPIRICAL EQUATIONS AND COMPRESSOR TEST DATA 95

PRESSURE RECOVERY FOR DIFFUSERS WITH DIFFERENT AREA RATIOS 96

COMPARISON OF EMPIRICAL EQUATIONS WITH EXPERIMENTAL RESULTS 9 7

v i i i

1. INTRODUCTION

Feas ib i l i t y demons t r a t ions o f son ic and nea r - son ic i n l e t s are numerous.

Table A i n t h e Appendix gives a b r i e f summary of tests tha t have been con-

duc ted dur ing the pas t few years . A l a r g e number of these were conducted

by a i r c r a f t and engine manufac turers wi th appl ica t ions d i rec ted toward

s p e c i f i c e n g i n e s . D e s p i t e t h i s c o n t i n u a l e f f o r t , a n d a l t h o u g h m o s t

demonstrat ion-of-concept s tudies yield impressive amounts of n o i s e

a t t e n u a t i o n , t h i s method of n o i s e r e d u c t i o n h a s n o t a c h i e v e d t h e s t a t u s

of a c c e p t a b i l i t y as other methods of noise reduct ion (i.e. a c o u s t i c

l i n e r s ) .

P r e s e n t l y , results f r o m l a b o r a t o r y o r f u l l s c a l e ( s t a t i c ) d e m o n s t r a t i o n s

h a v e n o t y e t b e e n t r a n s f e r r e d t o p r a c t i c a l a p p l i c a t i o n s on a i r c r a f t . A

de t a i l ed unde r s t and ing o f t he i n t e rac t ion o f t he acous t i c and aerodynamic

f i e l d s i s hampered by t h e f a c t t h a t t h e f l o w is near-sonic a t t h e t h r o a t ,

making bo th ana ly t ica l descr ip t ion and carefu l exper imenta l measurements

d i f f i c u l t due t o t h e dominance of non-linear effects. However, from a

compilat ion of t h e a v a i l a b l e d a t a , some genera l t rends can be d i scerned

which can be used t o i n d i c a t e c e r t a i n d e s i g n limits o r show i f some form

of boundary layer control is r equ i r ed .

The pu rpose o f t h i s r e sea rch p ro j ec t was t o d e v e l o p some fundamental aero-

dynamic and acous t i c i n fo rma t ion on t he son ic and nea r - son ic i n l e t s .

S .pec i f i ca l ly , t h i s r e sea rch p rog ram a t t empt s t o p rov ide ae rodynamic and

a c o u s t i c d a t a o f a s u f f i c i e n t l y g e n e r a l n a t u r e i n o r d e r t o k e e p o p e n t h e

o p t i o n s o f v a r i o u s c o n f i g u r a t i o n s i n choked i n l e t d e s i g n . The experimental

r e s e a r c h r e s u l t s r e p o r t e d h e r e are complemented by a c o n t i n u i n g t h e o r e t i c a l

study. The emphasis i n t h e e x p e r i m e n t a l s t u d y was on t h e t r a n s l a t i n g c e n t e r -

body type i n l e t , a l t hough compar i sons are made with f ixed geometry

(col lapsing cowl) inlets. With the except ion of one test on a Viper-8 j e t

e n g i n e i n 1966 (Cawthorn et a l . , see Appendix A ) , there have been no o ther

tests on t r a n s l a t i n g c e n t e r b o d y t y p e i n l e t s o u t s i d e t h o s e c o n d u c t e d b y o n e

of the authors (Lumsdaine, see Appendix A). Besides type comparison, the

s tudy i nvo lved an i nves t iga t ion of such parameters as area r a t i o , l e n g t h /

d i a m e t e r r a t i o , a n d l i p e f f e c t s t o d e t e r m i n e t h e i r i n f l u e n c e on t h e a c o u s t i c

and aerodynamic performance.

For purpose o f def in i t ion , sonic in le t s are those where the f low is

a c c e l e r a t e d u n t i l a s o n i c s u r f a c e e x i s t s n e a r t h e t h r o a t o f t h e i n l e t a n d

t h e i n l e t i s aerodynamically choked (or nearly so) wi th a corresponding

l a r g e n o i s e r e d u c t i o n . T h e s e i n l e t s are known t o b e v e r y e f f e c t i v e a c o u s t i -

c a l l y s i n c e t h e o n l y n o i s e t h a t c a n p r o p a g a t e u p s t r e a m i n t h i s case is t h e

noise escaping th rough the boundary l ayer . The p rob lems a s soc ia t ed w i th t he

s o n i c i n l e t are pr imar i ly aerodynamic , tha t is , how t o a c h i e v e a s o n i c

s u r f a c e w i t h t h e s h o r t e s t i n l e t , w i t h low d i s t o r t i o n , n e g l i g i b l e i n s t a -

b i l i t y , and minimum p r e s s u r e l o s s .

I n t h e n e a r - s o n i c ( o r a c c e l e r a t i n g ) i n l e t t h e f l o w n e a r t h e t h r o a t i s a t a

v e l o c i t y w h i c h r e s u l t s i n n o i s e r e d u c t i o n b u t h a s n o t y e t r e a c h e d t h e a e r o -

dynamic choking point. For one-dimensional flow, such a d i f f e r e n t a t i o n would

n o t b e n e c e s s a r y s i n c e t h e two types o f in le t s need on ly be d iv ided by an

a r b i t r a r y t h r o a t Mach number (say 0.8). This oversimplif icat ion, however ,

can cause problems when d a t a from d i f f e r e n t tests are compared, because for

i n l e t s of p r a c t i c a l l e n g t h and f o r n e a r - s o n i c v e l o c i t i e s , t h e f l o w d e v i a t e s

g r e a t l y from the one-dimensional potent ia l f low approximation. As can be

s e e n i n F i g u r e 1* (from Reference l ) , t h e c e n t e r l i n e t h r o a t Mach number i s

a ve ry i nadequa te pa rame te r , s ince i t r e m a i n s n e a r l y c o n s t a n t f o r t h e t h r e e

v e r y d i f f e r e n t f l o w c o n d i t i o n s ( a , b , c ) w i t h d i f f e r i n g a m o u n t s of mass flow.

T h i s c a n a l s o l e a d t o t h e m i s t a k e n n o t i o n t h a t n e a r s o n i c i n l e t s are e f f e c -

t i v e , when t h e t h r o a t c e n t e r l i n e Mach number ind ica tes subsonic f low even

though supersonic condi t ions exis t downstream of the throat .

*The f i g u r e s are given a t t h e end of each chapter .

2

Another point concerns the n o i s e a t t e n u a t i o n c h a r a c t e r i s t i c s . I n R e f e r e n c e 2 ,

tests were c o n d u c t e d u s i n g a n e j e c t o r w i t h i n l e t s o f d i f f e r e n t l e n g t h /

d i a m e t e r r a t i o b u t w i t h t h e same area r a t i o and noise source. The atterm-

a t i o n s o b t a i n e d h a v e d i f f e r e n t v a l u e s f o r t h e same m a x i m u m average axial

Mach number. (In t h i s r e p o r t , t h e a v e r a g e t h r o a t Mach number r e f e r s t o t h e

Mach number at the geomet r i c t h roa t . The maximum average axial Mach number

r e f e r s t o t h e maximum average Mach number in s ide t he channe l t aken no rma l

t o t h e f l o w d i r e c t i o n . ) It w a s noted i n t h e s e tests (see F i g u r e 2) t h a t

t h e s h o r t e r d i f f u s e r s t e n d t o b e more e f f e c t i v e i n terms o f n o i s e r e d u c t i o n

a t subsonic speeds, whereas the longer diffusers produce l i t t l e o r no n o i s e

r e d u c t i o n u n t i l t h e f l o w is close to being aerodynamical ly choked. The

r e a s o n f o r t h i s c a n b e a t t r i b u t e d e i t h e r t o a supersonic pocket near the

t h r o a t f o r t h e s h o r t e r d i f f u s e r s (due t o t h e smaller r ad ius o f cu rva tu re

a t the th roa t ) which is e f f e c t i v e i n r e d u c i n g some of the no ise p ropagated ,

o r t o t h e l a r g e r axial p r e s s u r e g r a d i e n t f o r t h e same o v e r a l l p r e s s u r e

d i f f e r e n c e . I n a n y case, F igu re 2 po in t s ou t t he i nadequacy of p l o t t i n g

no i se r educ t ion ve r sus a s i n g l e p a r a m e t e r ( i n t h i s case t h e maximum axial

Mach number) s ince ex t rapola t ion or compar ison of d a t a w i l l n o t b e r e l i a b l e

because condi t ions downstream of the throat have an inf luence on the noise

a t t e n u a t i o n . The percentage of maximum mass flow is a b e t t e r p a r a m e t e r f o r

p l o t t i n g n o i s e r e d u c t i o n . However, i t is very sens i t ive near choke f low; a

change of about 4 p e r c e n t i n mass flow changes the Mach number from 0.8

t o 1.0.

3

2. DESCRIPTION OF INLET CONFIGURATIONS

Table I g i v e s a summary o f t h e i n l e t s t e s t e d . F i g u r e s 3 t o 9 show the schemat ic

o f t h e s e i n l e t s i n c l u d i n g t h e area d i s t r i b u t i o n s at the cen te rbody pos i t i ons

t e s t e d . B e s i d e s t h e two b a s i c c o n f i g u r a t i o n s ( f i x e d a n d v a r i a b l e g e o m e t r y ) t h e

i n l e t s c a n b e c l a s s i f i e d i n t h e f o l l o w i n g way f o r test i d e n t i f i c a t i o n : I. Sonic

i n l e t f o r s u b s o n i c a n d s u p e r s o n i c r o t o r t i p s p e e d , 2. Accelerat ing (near-sonic)

i n l e t f o r s u b s o n i c t i p s p e e d , and 3 . A c c e l e r a t i n g i n l e t f o r s u p e r s o n i c t i p

speed . The ro tor ach ieved supersonic t ip speeds a t approximately 22,000 rpm.

However, f a r f i e l d d a t a d i d n o t show the p ropagat ion of mul t ip le pure tones

u n t i l f a n s p e e d was i n e x c e s s o f 23,500 rpm.

I n o r d e r t o o b t a i n t h e f l e x i b i l i t y r e q u i r e d t o d e t e r m i n e t h e d i f f e r e n t e f f e c t s

d e s i r e d i n t h i s s t u d y , t h e t r a n s l a t i n g c e n t e r b o d y t y p e i n l e t was used. This

t ype of i n l e t w a s a l s o f o u n d t o b e more s a t i s f a c t o r y i n earlier concept

e v a l u a t i o n s t u d i e s (as compared t o t h e c o l l a p s i n g c o w l , f o r e x a m p l e ) . The

p r e s e n t tests a l s o c o n f i r m t h e t r a n s l a t i n g c e n t e r b o d y t y p e t o b e more s u p e r i o r

i n r e g a r d t o p r e s s u r e l o s s , n o i s e a t t e n u a t i o n a n d f l o w s t a b i l i t y .

Configurat ion 1 was used t o de te rmine the e f fec t o f sonic and near sonic f low

on no i se fo r bo th subson ic and supe r son ic t i p speeds . Conf igu ra t ions 2 t o 4

were f i r s t u s e d t o d e t e r m i n e area r a t i o e f f e c t s on performance for sonic

i n l e t s w i t h s u b s o n i c t i p s p e e d s a n d later t o d e t e r m i n e l i p e f f e c t s f o r s u p e r -

s o n i c t i p s p e e d s . C o n f i g u r a t i o n s 5 and 6 were f ixed geomet ry t ype i n l e t s

mainly used for comparison purposes (aerodynamic and acoust ic) . The f i x e d

pos i t ions which were t e s t e d s i m u l a t e t h e c o l l a p s i n g c o w l t y p e i n l e t . Con-

f i g u r a t i o n s 7 t o 9 were low area r a t i o i n l e t s f o r t h e p r i m a r y p r u p o s e o f

t e s t i n g t h e i n f l u e n c e o f h i g h s u b s o n i c Mach number on s u p e r s o n i c t i p s p e e d

r o t o r n o i s e . The t h r e e b a s i c t y p e s of l i p s t e s t e d were the sho r t be l lmou th ,

t he s imu la t ed l and ing ( equa l t o 0.2 Mach number forward speed) and the

t a k e o f f l i p . The d i f f u s e r s e c t i o n s f o r C o n f i g u r a t i o n s 2 t o 4 and 7 t o 9 were n e a r l y i d e n t i c a l .

4

TABLE I - LIST OF INLET CONFIGURATIONS TESTED I

~~ ~ -

TYPE OF MAX. T I P SPEED LENGTH/ C.B. POS. AREA L I P (MAX. RPM) 1 DIAMETER I (CM) 1 REMARKS DESIGNED MAX.MASS 1 FLOW (KG/SEC)

CONF. NO.

1

2

3

ul

4

5

6

7

8

. 9

INLET TYPE

0.0 5 .1

12.7 17.8 20.3 25.4

- -

2.35 2.90 3.20 3.50

~~ ~

19.0

12 .o 9.75 8.6 7.9

- t r ans l a t ing centerbody

f l i g h t l i p

supersonic (25,000)

sonic (and near- s o n i c ) i n l e t 1.90

0.0 3.8 7.6

11.4

1.40 1.80 2.20 2.60

10.0 8.6 7.5 6.35

t r ans l a t ing centerbody

s h o r t bellmouth

subsonic (21,000)

sonic (and near- s o n i c ) i n l e t 1.68

~ ~~~ ~ ~


~ ~~ ~

simulated landing

f l i g h t l i p

simulated landing

- ~ ~~ ~ ~~~ ~ ~

0.0 1.40 3.8 1.80

11.4 2.60

0.0 1.40 3.8 1.80

11.4 2.60

~~

subsonic 1.68 (21,000) 2.20 7.6

subsonic 1.68 (21,000) 2.20 7.6

subsonic (20,000)

1.54 2.60 -

~~~ ~~ ~~ -~ ~

10.0 sonic (and near-

7.5 s o n i c ) i n l e t 8.6

6.35

10.0 sonic (and near-

7.5 s o n i c ) i n l e t 8.6

6.35

sonic (and near- s o n i c ) i n l e t

6.35


f ixed geometry

~ ~~~ -

f ixed geometry

~~~ ~~ ~~ ~ ~

sonic (and near- s o n i c ) i n l e t 1 7.5 landing 1 1.54 1 - 1 2.30

f ixed geometry

f ixed geometry

f ixed geometry

s h o r t supersonic bellmouth (25,000)

1.68 1.20 - near sonic i n le t 15.2

landing 1 1.68 1 - near sonic i n l e t 1 15.2

I I I I I

I I 1

near sonic i n l e t 1 15.2 I

3. PROCEDURE FOR INLET DESIGN

Two t y p e s o f i n l e t s were des igned fo r t hese tests: 1) wi th a t r a n s l a t i n g

centerbody and 2) without centerbody. The t r ans l a t ing cen te rbody t ype

i n l e t s were designed to produce a h i g h p r e s s u r e g r a d i e n t n e a r t h e t h r o a t ;

t h e i n l e t s w i t h o u t c e n t e r b o d y were designed with a more g r a d u a l p r e s s u r e

g r a d i e n t f o r t h e same o v e r a l l area r a t i o . The s h a r p p r e s s u r e g r a d i e n t was

p l a c e d n e a r t h e t h r o a t ( w h e r e f o r t r a n s o n i c f l o w t h e s t r e a m l i n e s are v e r y

s t i f f ) b e c a u s e earlier tests conducted under this program showed t h a t f o r

a g i v e n o v e r a l l p r e s s u r e r a t i o , s o u n d a t t e n u a t i o n a t n e a r s o n i c c o n d i t i o n s

was more e f f e c t i v e when t h e p r e s s u r e g r a d i e n t was h igh . This is a l s o e v i d e n t

from Figure 2. Recent tests by General Electric3 c o n f i r m e d t h i s e f f e c t .

Because a h igh cons t an t Mach number s e c t i o n ( c o n s t a n t a r e a ) is a l s o e f f e c -

t ive i n r e d u c i n g t h e n o i s e , p a r t i c u l a r l y f o r s u p e r s o n i c t i p s p e e d s , a r a p i d

d i f f u s i o n n e a r t h e t h r o a t was followed by a c o n s t a n t area s e c t i o n i n t h e

design.

The l e n g t h / d i a m e t e r r a t i o f o r t h e area r a t i o s of t h e p r e l i m i n a r y d e s i g n was

se l ec t ed based on some earlier work which found t h a t a n optimum l e n g t h /

d i a m e t e r r a t i o e x i s t e d f o r a given area r a t i o . F i g u r e 10 shows t h e r e s u l t s

of some earlier tests conducted under this program using six annular-type

i n l e t s of d i f f e r e n t d i f f u s e r l e n g t h s b u t w i t h t h e same area r a t i o o f 3 . 2 .

Although parameters such as throat blockage due to boundary layer develop-

ment and increasing Mach number have a c o n s i d e r a b l e e f f e c t on dec reas ing

the p re s su re r ecove ry , this does not seem to have a s i g n i f i c a n t i n f l u e n c e

on the optimum p o i n t . Also, a l i t e r a t u r e s e a r c h u n c o v e r e d a l a r g e number of

tests w i t h c i r c u l a r and s q u a r e d i f f u s e r s ; some of the impor tan t resu l t s are

compiled i n F i g u r e 11 which shows the approximate re la t ionship be tween area

r a t i o and optimum leng th /d i ame te r r a t io . Fo r t he ca se o f , annu la r d i f fuse r s

t h e number of tests tha t cou ld be u sed fo r such a graph were n o t as numerous;

6

t hey are b r o u g h t t o g e t h e r i n F i g u r e 12. The term "f low d iameter" for an

annu la r d i f fuse r i nd ica t e s t he i nne r d i ame te r minus t he cen te rbody

diameter a t t h e exit. From these g raphs i t can be seen that the cen terbody

t y p e i n l e t r e q u i r e s a s h o r t e r l e n g t h f o r t h e same area r a t i o t h a n a n i n l e t

without centerbody.

From a n i n i t i a l estimate o f t h e l e n g t h / d i a m e t e r r a t i o f o r a given area r a t i o

requi rement , the in le t w a s designed based on a smooth area p r o g r e s s i o n f o r

bo th t he choked and unchoked posit ions. After the contour w a s l a i d o u t , a

t r anson ic po ten t i a l f l ow p rogram4 was u s e d t o c a l c u l a t e t h e f l o w f i e l d . The

r e s u l t s were t h e n s u b s t i t u t e d i n t o a boundary layer program to determine

co r rec t ions fo r t he con tour . In a lmos t a l l t h e s e c a l c u l a t i o n s , some f low

s e p a r a t i o n was p r e d i c t e d by the boundary l ayer ana lys i s because o f the wide

range of o p e r a t i n g c o n d i t i o n s o c c u r i n g i n t h e i n l e t and i n o r d e r t o k e e p

t h e i n l e t w i t h i n r e a s o n a b l e l e n g t h f o r a given area r a t i o . F i g u r e s 13 t o 17

g ive t he s t r eaml ines and Mach number d i s t r i b u t i o n a l o n g t h e wa l l and center-

body fo r Conf igu ra t ion 1 ( t a k e o f f c o n f i g u r a t i o n o r f l i g h t l i p ) f o r f i v e of t h e

cen te rbody pos i t i ons t e s t ed . These r e su l t s were used in t he boundary l aye r

program to determine boundary layer growth. Because of the high adverse

p r e s s u r e g r a d i e n t when the centerbody is t r a n s l a t e d t o 12.7 cm and beyond,

some f low separat ion had been predicted f rom boundary layer calculat ions.

However, f o r t h i s p a r t i c u l a r c o n f i g u r a t i o n , t h e a c t u a l s e p a r a t i o n was more

severe as i n d i c a t e d b y t h e e x p e r i m e n t a l s u r f a c e p r e s s u r e d i s t r i b u t i o n a n d

t h e t o t a l p r e s s u r e p r o b e s a t t h e e x i t . The exper imenta l sur face Mach numbers

are a l s o i n d i c a t e d o n t h e s e f i g u r e s .

Three t y p e s o f l i p s were d e s i g n e d : t h e f l i g h t l i p , t h e c o n t o u r e d l i p simu-

l a t i n g t h e s t r e a m l i n e a t 0.2 f r ees t r eam Mach number, and a shor t be l lmouth .

Furthermore, the f ixed geometry inlets were designed with a s t r a i g h t w a l l

s i n c e c o n t o u r i n g t h e wall does not have a s i g n i f i c a n t i n f l u e n c e o n t h e a e r o -

dynamic performance. These i n l e t s were des igned wi th a s l i g h t l y s h o r t e r

l eng th t han t he optimum ( i f F i g u r e 11 is fol lowed) for the purpose of main-

t a i n i n g t h e same area r a t i o and approximately the same leng th /d i ame te r r a t io

as t h e t r a n s l a t i n g c e n t e r b o d y t y p e i n l e t s (see Table I) i n o r d e r t o compare

t h e e f f e c t i v e n e s s of t h e two types of i n l e t s .

7

" .

4. TEST FACILITY, INSTRUMENTATION AND PROCEDURE

The i n l e t s were tes ted in the anechoic-chamber t ransonic compressor fac i l i ty

of t h e NASA-Langley Research Center. The test v e h i c l e was a 30.5 cm-tip

d iameter s ing le-s tage t ransonic compressor wi th 19 r o t o r b l a d e s . F i g u r e 18

shows the anechoic chamber fac i l i ty wi th one o f the choked in le t s in p lace .

The f a r f i e l d n o i s e measurements were made w i t h a movable boom microphone

(foreground) a t 4.57 m f r o m t h e i n l e t .

F igu re 1 9 shows the i n s t rumen ta t ion room where the speed of the compressor,

t h e p o s i t i o n of t he t r ans l a t ing cen te rbody , and t he boom microphone posi t ion

were c o n t r o l l e d . I n a d d i t i o n , a e r o d y n a m i c o u t p u t w a s recorded on punch cards

f rom the scaniva lves , and acous t ic ou tput was recorded on magnetic tape. On-

l i n e f a r f i e l d n o i s e d a t a were taken wi th a 1/3-octave real-time ana lyze r a t

two l o c a t i o n s by an independent system for comparison and quick-look monitor-

ing. The fol lowing aerodynamic instrumentat ion was used:

1. Four s t a t i c p r e s s u r e t a p s were l o c a t e d c i r c u m f e r e n t i a l l y u p s t r e a m o f

t h e cowl h i g h l i g h t .

2. S i x t y s t a t i c p r e s s u r e t a p s l o c a t e d a x i a l l y i n two c i r c u m f e r e n t i a l

s ta t ions (12 on one and 48 on the o ther ) were used for Configura-

t i o n 1; 45 were used for each of t h e o t h e r c o n f i g u r a t i o n s . I n

a d d i t i o n , t h e r e were 12 s t a t i c pressyre t aps on each cen terbody.

3. One k i e l h e a d t o t a l p r e s s u r e traverse probe and two 20-element t o t a l

p r e s s u r e r a k e s were l o c a t e d a t t h e ex i t p l a n e of t h e i n l e t .

4 . One f ixed k i e l head p robe r ead p re s su re and t empera tu re a t t h e e x i t

plane.

5. Cen te rbody t r ans l a t ion was cont ro l led by a h y d r a u l i c c y l i n d e r . The

centerbody pos i t ion was indica ted by a vo l tme te r a t t h e c o n t r o l

panel .

8

A schematic of some of the aerodynamic instrumentat ion and their approximate

l o c a t i o n is shown on F igure 20 f o r a t r a n s l a t i n g c e n t e r b o d y t y p e i n l e t and

on Figure 2 1 f o r a f ixed geomet ry t ype i n l e t .

The acous t i c i n s t rumen ta t ion cons i s t ed o f

1. A travelingboom 0.635-cm condenser microphone located '4.57 m from

t h e i n l e t p l a n e and reading a t 15-degree increments,

2. 0.635-cm condenser microphones also located a t 4.57 m fo r on - l ine

d a t a a c q u i s i t i o n ,

3 . an acous t ic t ravers ing probe p laced a t t h e exit p l a n e o f t h e i n l e t ,

and

4. four flush-mounted 0.635 cm microphones (or six depending on t h e

i n l e t ) l o c a t e d on the cowl as shown on F igu res 3 t o 9.

A l l microphones were c a l i b r a t e d b e f o r e a n d a f t e r e a c h test . The boom micro-

phone measu red t he f a r f i e ld no i se a t 15-degree intervals f rom O-degrees to

90 degrees . Two f a r f i e l d m i c r o p h o n e s a t O-degrees and 15-degrees recorded

t h e f a r f i e l d d a t a c o n t i n u o u s l y a n d were monitored on a 1/3-octave real-time

analyzer . This ana lyzer was also used to monitor choking speed and to regu-

l a te the ope ra t ing po in t s fo r t he compresso r . Acous t i c traverses were made

only near the choke po in t and t raverse da ta were recorded for a minimum of

10 s e c o n d s ( f o r s t a b i l i z a t i o n ) a t each po in t . S teady acous t ic da ta (on wal ls ,

f a r f i e l d , e t c . ) were recorded for a minimum of one minute a t each da ta

p o i n t . Only one traverse probe was operated a t a given t i m e . The a c o u s t i c

traverse was immersed a t f i v e r a d i a l p o s i t i o n s f o r d a t a a c q u i s i t i o n .

Aerodynamic in l e t pe r fo rmance da t a were recorded on punch cards, with each

parameter sampled several times p e r d a t a p o i n t . The t o t a l p r e s s u r e k i e l p r o b e

d a t a w a s recorded on an X-Y p l o t t e r . Data were recorded a t fixed immersion

p o i n t s and cont inuous ly dur ing the traverse; each f ixed immersion was sampled

f o r 10 seconds.

T y p i c a l l y , t h e test o p e r a t i o n f o r a t r a n s l a t i n g c e n t e r b o d y i n l e t f o l l o w e d

this sequence: With the centerbody completely re t racted, the compressor rotor

9

w a s g r a d u a l l y a c c e l e r a t e d t o some p r e s c r i b e d maximum rpm a n d s t a b i l i z e d ;

a c o u s t i c and aerodynamic data were recorded a t p r e s c r i b e d i n t e r v a l s d u r i n g

t h i s a c c e l e r a t i o n , and the centerbody was t h e n t r a n s l a t e d t o t h e n e x t test

p o s i t i o n .

The on-line 1/3-octave real-time analyzer monitored the sound levels u n t i l

the choking (acous t ic ) rpm p o i n t was reached. Data were reco rded fo r several

rpm and centerbody posi t ions. The b a c k p r e s s u r e r a t i o was a d j u s t e d t o n e a r

m a x i m u m ( b l a d e s t a l l ) and a l s o t o minimum (valve wide open) with one inter-

media te po in t t o g i v e d i f f e r e n t b l a d e l o a d i n g f o r t h e same rpm. A t y p i c a l

s e t of d a t a p o i n t s f o r a g iven cen te rbody pos i t i on i s shown i n F i g u r e 22 f o r

i n l e t C o n f i g u r a t i o n 1. O t h e r i n l e t s were run w i th t he back p re s su re va lve

wide open only.

10

5 . DISCUSSION OF RESULTS

Noise a t t enua t ion fo r i n l e t s unde r nea r - son ic cond i t ions seems t o depend on

the source ( t ip speed , peak f requency) as w e l l as on the f low condi t ions

(axial v e l o c i t y g r a d i e n t , r a d i a l g r a d i e n t , p e a k Mach number); thus the choice

o f pa rame te r s t o r ep resen t t he acous t i c and aerodynamic conditions should be

made v e r y c a r e f u l l y . F i g u r e s 23 t o 27 g i v e t h e f i v e p a r a m e t e r s which l a r g e l y

def ine the per formance of ln le t Conf igura t ion 1: a) n o i s e a t t e n u a t i o n ,

b) mass flow, c) average m a x i m u m Mach number (cowl and centerbody), d) pressure

recovery, and e) p r e s s u r e d i s t o r t i o n . The d e t a i l e d Mach number d i s t r i b u t i o n s

along the cowl and centerbody for each one of t h e s e d a t a p o i n t s are g iven i n

a s e p a r a t e volume (present ly being compiled) . The c o n s t a n t n o i s e a t t e n u a t i o n

l i n e s o n t h e mass f low versus rpm graphs ( i .e. F igu re 2 4 ) are i n t e r p o l a t i o n s .

F o r t h i s i n l e t o n l y , t h e n o i s e a t t e n u a t i o n v a l u e s were taken wi th respec t

t o t h e f u l l y r e t r a c t e d p o s i t i o n . F o r a l l o t h e r c o n f i g u r a t i o n s , t h e v a l u e s

were t aken w i th r e spec t t o a basel ine, normally without the centerbody. It

is t y p i c a l of high area r a t i o s h o r t i n l e t s t h a t t h e n o i s e a t t e n u a t i o n is

v e r y g r a d u a l , t h a t is, the no i se does no t d rop r ap id ly as the choke point

is approached. This was a l s o e v i d e n t i n p a s t tests u s i n g a n e j e c t o r . I n

F igu re 2 ( f o r t h e same area r a t i o b u t d i f f e r e n t l e n g t h s ) , t h e n o i s e r e d u c t i o n

as a f u n c t i o n of rpm is g r a d u a l f o r t h e s h o r t e r i n l e t b u t i n c r e a s e s s h a r p l y

f o r t h e l o n g i n l e t . The d i s t o r t i o n l e v e l f o r t h i s h i g h area r a t i o s h o r t i n l e t

is very low, due t o a long cons tan t area section downstream of t h e t h r o a t .

F igu res 28 t o 42 are pe r fo rmance r e su l t s fo r i n l e t Conf igu ra t ions 2 t o 4 .

These conf igu ra t ions d i f f e r on ly i n t he shape o f t he l i p ; t he cen te rbody

a n d t h e d i f f u s e r s are t h e same. F o r t h e s e i n l e t s w i t h r e l a t i v e l y low area

r a t i o s , n o i s e a t t e n u a t i o n is qui te sudden and occurs wi th in 5 t o 10 pe rcen t

of maximum mass flow as can be seen f rom F igures 43 to 45 w h e r e t h e f a r

11

f i e l d n o i s e is p l o t t e d as a func t ion of mass f low. The noise reduct ion

o c c u r s f o r a l l t h r e e c o n f i g u r a t i o n s i n t h e i n t e r v a l o f 90 t o 100 pe rcen t

of mass flow. It should be no ted tha t the major no ise d rop is w i t h i n a

narrow ( 5 percent ) range of mass flow. On a one-d imens iona l bas i s , th i s

means a move from very l i t t l e n o i s e r e d u c t i o n t o a l m o s t f u l l c h o k e f o r

a 5 percent change i n area. Near Mach one, a 5 p e r c e n t c h a n g e i n area

causes the average Mach number to change from 0.75 t o 1.0. Also, when

the cen terbody w a s t r a n s l a t e d , t h e r e was a n i n c r e a s e i n t h e n o i s e level

from the compressor . This can be readi ly seen in Figures 28, 33 , and 38.

A s a r e s u l t of mechanical problems i n t h e d e s i g n , a s l i g h t l y h i g h e r

p r e s s u r e l o s s than expected w a s e n c o u n t e r e d w i t h t h e t r a n s l a t i n g t y p e

i n l e t because t he re was a d iscont inui ty be tween the cy l inder and cen ter -

body (which s l i d o v e r t h e c y l i n d e r ) . Also , t h e r e was a groove which was

exposed when the cen terbody was t rans la ted forward . This g roove was

n e c e s s a r y i n o r d e r t o k e e p t h e c e n t e r b o d y f r o m r o t a t i n g , b u t when a

c o n f i g u r a t i o n was tes ted wi th the cen terbody, the p resence o f the g roove

may h a v e c o n t r i b u t e d t o t h e i n c r e a s e d f a r f i e l d n o i s e . The n o i s e l e v e l

r o s e as the cen terbody was t r a n s l a t e d f u r t h e r , e x p o s i n g a l a r g e r segment

of t he g roove ( i . e . see F igure 4 5 ) . Also, w i th t he cen te rbody t r ans l a t ed ,

t he h ighe r p re s su re g rad ien t caused some f low separa t ion which cont r ibu ted

t o t h e i n c r e a s e d n o i s e l e v e l .

F igu res 46 t o 50 show t h e p e r f o r m a n c e r e s u l t s f o r two fixed-geometry type

i n l e t s ( s i m u l a t i n g a c o l l a p s i n g cowl type i n l e t ) w i thou t cen te rbody . A l -

though the area r a t i o s are n o t v e r y h i g h (see Table I), b o t h i n l e t s p e r -

formed very poorly. The flow was so uns t ab le . - and t he mon i to r fo r t he f a r

f i e l d n o i s e showed s u c h l a r g e v a r i a t i o n s t h a t t h e a v e r a g e n o i s e r e d u c t i o n

shown on Figure 46 is very approximate. I t appeared that even under choke

c o n d i t i o n s t h e amount of n o i s e r e d u c t i o n was ve ry small. The test w a s

discont inued a t around 17,000 rpm because of severe p res su re and acous t i c

f l u c t u a t i o n s f o r i n l e t C o n f i g u r a t i o n 5 , and a t 19,000 rpm for Conf igura-

t i o n 6 . The h i g h d i s t o r t i o n c a n b e s e e n o n F i g u r e 50 . N o t e t h a t t h e Mach

number p l o t ( F i g u r e 48) is rep resen ted by t h e wall Mach number and n o t t h e

average Mach number as in t he p rev ious g raphs . The average values would be

12

lower than the peak wall Mach number. F igure 51 is a graph of noise reduct ion

as a func t ion of mass f low fo r Conf igu ra t ions 5 and 6. These poor r e s u l t s

were qu i t e unexpec ted , s ince t he same g u i d e l i n e s were f o l l o w e d i n t h e d e s i g n

of Configurat ions 5 and 6 as f o r t h e o t h e r i n l e t s .

Configurat ions 7 t o 9 were r u n t o o b t a i n b a s e l i n e d a t a f o r C o n f i g u r a t i o n s

2 t o 4 as w e l l as d a t a f o r n o i s e r e d u c t i o n a t h i g h t h r o a t Mach numbers

when t h e f a n is a t supe r son ic t i p speeds . They were also used t o d e t e r m i n e

t h e l i p e f f e c t . F i g u r e s 52 t o 55 show t h e b a s i c f l o w c h a r a c t e r i s t i c s o f

Configurat ions 7 t o 9. Note t h a t h e r e , t o o , t h e Mach number p l o t i s re-

presented by t h e wal l Mach number and no t the average Mach number. F igu re 56

shows t h e f a r f i e l d a c o u s t i c c h a r a c t e r i s t i c s f o r t h e s e t h r e e i n l e t s ( w i t h

Configurat ion 1 a t 0 c m centerbody displacement given for comparison). For

lower rpm (lower Mach number) t he i n l e t w i th t he cen te rbody appea r s t o

g i v e a few decibels more a t tenuat ion than the same i n l e t w i t h o u t c e n t e r -

body. Th i s i s o n l y t r u e when comparisons are made wi th the cen terbody re-

t racted. Al though Configurat ion 9 ( f l i g h t l i p ) h a s a c o n s i s t e n t l y h i g h e r

n o i s e l e v e l a t lower rpm t h i s i s decreasing more rapidly a t higher rpm.

T h i s i s more ev iden t f rom the r e su l t s of blade passage f requency shown i n

F igu re 57. Thus the h igh ve loc i ty a round t he l i p seems t o i n f l u e n c e t h e

a t t e n u a t i o n .

F igu re 58 shows the f requency spec t rum for in le t Conf igura t ion .2 a t approxi-

mately 17,500 rpm (an exac t rpm is d i f f i c u l t t o m a i n t a i n ) and d i f f e r e n t

centerbody posi t ions. It can be s een t ha t t he amount of n o i s e r e d u c t i o n a t

the blade passage f requency is much h ighe r t han t he ove ra l l no i se r educ t ion

and t h a t t h e h i g h e r f r e q u e n c i e s are a t t e n u a t e d much more than the lower

f r e q u e n c i e s f o r h i g h Mach numbers. A t f u l l y choked conditions, however, the

amount of n o i s e r e d u c t i o n of t h e d i f f e r e n t f r e q u e n c i e s is more uniform, as

s e e n i n F i g u r e 59.

1 3

6. COMPARISON OF RESULTS

Configurat ion 1 was u s e d t o test choking a t ve ry h igh area r a t i o s (maximum

of 3.5) whi le main ta in ing a reasonable l ength /d iameter ra t io o f 1 .9 . The

remain ing conf igura t ions were shor te r wi th lower area ra t io s . These con-

f i g u r a t i o n s showed some f low sepa ra t ion when the cen terbody was n e a r

maximum t r a n s l a t i o n . The computer program r e q u i r e d t h e i n t r o d u c t i o n o f a

forward speed ( f ree stream Mach number = 0.1) which i s accep tab le du r ing

f l i g h t b u t d o e s n o t r e p r e s e n t t h e s t a t i c t es t c o n d i t i o n s a c c u r a t e l y . The

a c c e l e r a t i o n a r o u n d t h e l i p i n C o n f i g u r a t i o n 1 d u r i n g s t a t i c tests caused

some separa t ion and i s p a r t i a l l y r e s p o n s i b l e f o r t h e low pressure recovery .

I n comparing Configurations 2 t o 4 t o de t e rmine t he i n f luence o f t he l i p ,

F igu re 60 shows t h e d i f f e r e n c e i n n o i s e r e d u c t i o n f o r t h e t h r e e l i p s w i t h

the cen te rbody r e t r ac t ed . It i s s e e n t h a t a t lower rpm t h e r e is some n o i s e

r e d u c t i o n f o r t h e f l i g h t l i p b u t n o t f o r t h e c o n t o u r e d l i p s . The acce le ra -

t i on a round t he l i p causes a high-veloci ty pocket around the l ip which

improves the acoust ic performance, but decreases pressure recovery.

F igu res 61 t o 64 show t h e i n f l u e n c e of t h e l i p on no i se a t t enua t ion and

pressure recovery a t d i f f e r e n t p o s i t i o n s of the centerbody. The s h o r t b e l l -

m o u t h a p p e a r s t o h a v e b e t t e r n o i s e r e d u c t i o n c h a r a c t e r i s t i c s t h a n t h e f l i g h t

l i p . C o n t o u r i n g t h e l i p s l i g h t l y ( t o s i m u l a t e a p p r o a c h s t r e a m l i n e p a t t e r n s )

does not seem t o i m p r o v e t h e n o i s e r e d u c t i o n c h a r a c t e r i s t i c s . I n f a c t , a t

some h ighe r mass flow rates, t h e f l i g h t l i p a p p e a r s t o h a v e a h ighe r no i se

r e d u c t i o n f o r t h e same p r e s s u r e loss compared to the s imula ted l anding l i p .

For small area r a t i o s , t h e p r e s s u r e r e c o v e r y i s high enough s o that changes

i n t h e l i p do not have a l a r g e e f f e c t on the p re s su re r ecove ry . The d i f f e r -

ence i n p r e s s u r e r e c o v e r y seems t o i n c r e a s e f o r i n c r e a s i n g area r a t i o s u n t i l

a f a i r l y l a r g e area r a t i o is reached. With forward speed, the flow w i l l more

14

c lose ly approximate the shor t be l lmouth , and thus the no ise reduct ion and

pressure recovery should improve over s t a t i c cond i t ions . Tests in Re fe rence 5

showed that dur ing choke and with 100 fee t / sec ups t ream b lowing the p ressure

r ecove ry i nc reased on ly s l i gh t ly . However, t h e area r a t i o f o r these inlets

was l a r g e ; t h u s it a p p e a r s f o r h i g h area ra t io , fo rward speed may n o t i n -

crease t h e p r e s s u r e r e c o v e r y a n d n o i s e a t t e n u a t i o n b u t s h o u l d d o so f o r low

area r a t i o . T h i s is a l s o c o n s i s t e n t with t h e r e s u l t s of Figures 6 1 t o 64.

When F igures 61, 62 and 6 3 are compared wi th F igu re 64 ( l a r g e area r a t i o ) ,

i t i s s e e n t h a t t h e r e is ve ry l i t t l e d i f fe rence be tween the s imula ted l anding

l i p and t h e s h o r t b e l l m o u t h .

F igu re 65 compares t h e l i p e f f e c t w i t h and without centerbody for two con-

f i g u r a t i o n s ( s h o r t b e l l m o u t h a n d f l i g h t l i p ) . C o n s i d e r i n g t h e c a s e w i t h o u t

centerbody, i t i s q u i t e clear t h a t n o i s e r e d u c t i o n is more g r a d u a l f o r t h e

case of t h e f l i g h t l i p ( C o n f i g u r a t i o n 4 ) t h a n f o r t h e s h o r t b e l l m o u t h . The

noise reduct ion for the shor t be l lmouth occurs qu i te suddenly and over a

ve ry small span of increasing rpm. This also p o i n t s o u t t h e d i f f i c u l t y of

s imulat ing forward speed using s t a t i c tests. F igu re 66 i s a p o l a r p l o t of

t h e f a r f i e l d n o i s e a t 15 -degree i n t e rva l s fo r t he f i xed geomet ry i n l e t s

(Configurat ions 7 , 8 and 9 ) . The d i f f e r e n c e i n t h e l i p s h a p e c a u s e s a

s l i g h t d i p a t an angle of 15 d e g r e e s f o r t h e case of t he con toured l i p ;

f o r t h e f l i g h t l i p , t h e d i s t r i b u t i o n i s qu i t e un i fo rm.

Figure 67 shows t h e i n f l u e n c e o f a change i n compresso r p re s su re r a t io on

n o i s e . V a r i a t i o n i n t h e b a c k p r e s s u r e of the compressor does not seem t o

c h a n g e t h e n o i s e i n t h e f a r f i e l d , e x c e p t when the back pressure i s high

enough t o c a u s e t h e i n l e t t o go from choke point t o unchoke; however, t h e

mass f low a l so d rops . A s s e e n f r o m t h i s f i g u r e , t h e r e i s ve ry l i t t l e change

i n t h e f a r f i e l d n o i s e as a r e s u l t of changing the compressor back pressure

f o r t h e same mass f low. The data points were s l i g h t l y s c a t t e r e d ; t h u s t h e

s t r a i g h t l i n e s are only approximate. The influence of compressor blade

loading on far f i e l d n o i s e w i t h t h e i n l e t a t h igh Mach numbers i s an im-

portant problem, and fur ther tests are n e c e s s a r y t o g a i n more prec ise and

u s e f u l i n f o r m a t i o n i n t h i s area.

3.5

Figure 68 is a . compar i son o f t he r e su l t s of t h e p r e s e n t tests w i t h d a t a

publ i shed in Reference 3 u s i n g t h e same compressor running a t supe r son ic

relative Mach numbers a t t h e t i p . It w a s f o u n d t h a t t h e amount of n o i s e

r e d u c t i o n was much less t h a n t h a t p r e d i c t e d b y t h e m e t h o d s o f e i t h e r

Fisher6 or Matthews . 7

The c ruc ia l ae rodynamic pa rame te r s fo r any i n l e t t o be u sed on a n a i r c r a f t

engine are pressure recovery and d i s tor t ion , and low pressure recovery i s

usua l ly coup led w i th h igh d i s to r t ion . The cont ro l l ing geometr ic parameters

are length /d iameter ra t io and area r a t i o . The area r a t i o r e q u i r e d i s d ic -

t a t e d by a g i v e n a p p l i c a t i o n , a n d t h e optimum leng th /d i ame te r r a t io is

se l ec t ed f rom wh ich t he bes t des ign con tour is made. It a p p e a r s t h a t l a r g e

area r a t i o choked i n l e t s would be d i f f i c u l t t o d e s i g n w i t h r ea l i s t i c

lengths . For example, wi thout modif icat ion to the JT3D e n g i n e , i f i t were

necessa ry t o des ign t he i n l e t t o choke (o r nea r ly choke ) du r ing approach

€ o r a Boeing 707 wi th no rma l l oad and fu l l f l aps , an area r a t i o of h ighe r

t han t h ree would be r equ i r ed . But f o r a STOL a i r c r a f t w i t h blown f l a p s ,

t h e r e q u i r e d area r a t i o would b e less than two. The inf luence o f area

r a t i o on aerodynamic and acoustic performance thus was one of the parameters

s tud ied , and F igure 69 g i v e s a comparison of i n l e t s w i t h d i f f e r e n t area

r a t i o s and length /d iameter ra t ios . It a p p e a r s h e r e t h a t t h e l e n g t h / d i a m e t e r

r a t i o is no t as c r u c i a l as t h e area r a t i o . F o r l a r g e area r a t i o s , t h e choked

i n l e t seems to p roduce h igh losses and d i s tor t ion even wi th a centerbody-

type i n l e t . Fo r t he cen te rbody- type i n l e t w i th l ow area r a t i o s (A exit / ' throat less than 2 ) , t he choked i n l e t ope ra t e s w i th accep tab le p re s su re r ecove ry and

d is tor t ion and f rom a l l ind ica t ions appea r s t g provide s tab le f low.

.

16

7 . EMPIRICAL EQUATIONS TO ESTIMATE NOISE ATTENUATION

Assuming the forward and rearward no ise d i s t r ibu t ion f rom the fan to be

equa l and assuming t h a t t h e mean f l o w v e l o c i t y i s one-dimensional and

constant , the sound propagat ion f rom a cy l ind r i ca l e l emen t dA is , i n t h e

fo rward d i r ec t ion ,

dWf = (1 - M)T = (1 - M) 2 r de d r dW I

and rearward

dW,, = (1 + M)F = (1 + M) 7 r de d r dW I

where I is the sound i n t ens i ty , M t h e Mach number, r t h e r a d i a l

p o s i t i o n , W f , Wr the total forward and rearward propagating sound power,

and 8 t h e a n g u l a r p o s i t i o n .

When Equations (1) and (2) are in t eg ra t ed ove r a c i r c u l a r p l a n e , t h e

f o l l o w i n g c o r r e c t i o n f a c t o r s are obtained:

1 - 1 1"' '1: l + M Cf - - = -

Thus the no ise reduct ion in the forward d i rec t ion due to the oppos ing f low

is

AdB = -10 log Cf = -10 log[-] 1 l - M (3)

This empi r i ca l fo rmula has been u sed t o estimate n o i s e r e d u c t i o n i n h i g h

speed f low ' "O. A modi f i ca t ion of t h i s f o r m u l a ( a g a i n u s i n g t h e mean

t h r o a t Mach number as t h e v a r i a b l e ) was suggested i n Reference 8. The

17

problem wi th us ing these formulas is the l ack o f a p r o p e r d e f i n i t i o n f o r

the Mach number of an actual i n l e t a n d is q u i t e e v i d e n t h o m F i g u r e 1.

It is more real is t ic to desc r ibe no i se r educ t ion w i th an empi r i ca l fo rmula

based on percent o f maximum mass flow. One such formula was r e c e n t l y g i v e n

in Re fe rence 11: for one-d imens iona l i sen t ropic f low of a i r , t h e r e l a t i o n -

ship between Mach number and percent mass f low is

Th/imax = 1.73M(1+0.2M2)-3

Equation (3) can be r ewr i t t en i n t he fo rm

AdB = -10 l o g [ 1 1 - f ““/imax> 1

To de termine f ( i /kax) f rom exper imenta l da ta , i t was found tha t a

func t ion such as where 0 < a 5 1, a p p e a r s t o b e s t f i t t h e

exper imenta l da ta . The cons t an t c1 appears to depend on t h e p r e s s u r e

recovery near choke point . Figure 70 shows t h e v a l u e of a as a f u n c t i o n

of pressure recovery. The probable cause of the re la t ionship between a and pressure recovery is t h a t f o r i n l e t s w i t h r e l a t i v e l y low p r e s s u r e re-

c o v e r y , t h e r e e x i s t s a r a the r t h i ck boundary l aye r where t he no i se can

escape. This is the reason why t h e r e e x i s t s a l a r g e v a r i a t i o n i n t h e

r e p o r t e d r e s u l t s of t h e e f f e c t i v e n e s s of choked i n l e t s ( f r o m a few dec ibe ls

as given by Cawthorn e t a1 t o o v e r 40 dB i n some references such as 5,8,14) .

I n l e t s w i t h h i g h l o s s e s w i l l no t p roduce la rge no ise reduct ions even though

the f low is choked. However, h i g h l o s s i n l e t s n o r m a l l y p r o d u c e l a r g e r n o i s e

r e d u c t i o n t h a n i n l e t s w i t h low loss a t subsonic Mach numbers, possibly

because of t h e l a r g e g r a d i e n t s e x i s t i n g i n t h e s e i n l e t s . T h i s phenomenon is

now t h e s u b j e c t o f f u r t h e r i n v e s t i g a t i o n .

Figures 7 1 and 72 are comparisons of Equation (5) wi th expe r imen ta l da t a

r e p o r t e d i n R e f e r e n c e s 1 4 and 15. Reference 14 used the e jector as a source,

and Reference 16 used a 12-inch compressor. Also plotted for comparison i s

18

Equation ( 3 ) . It is n o t n e c e s s a r y t o compare, Equation (5) w i t h t h e p r e s e n t

r e s u l t s s i n c e the c u r v e f o r a ver sus p re s su re r ecove ry was de r ived by

u s i n g the present exper imenta l da ta ; obvious ly the empir ica l formula agrees

v e r y w e l l w i th t he expe r imen ta l r e su l t s .

However, i n o r d e r t o g a i n some conf idence in the method, the p ressure

r ecove ry fo r two t y p e s o f d i f f u s e r s was c a l c u l a t e d - f o r d i f f e r e n t area

r a t i o s and a t two mass f l o w r a t i o s o f i n t e r e s t . The r e s u l t s are shown on

Figure 73. The lower curves are f o r l a r g e area r a t i o d i f f u s e r s , and t h e

upper curves are f o r s t r a i g h t wall d i f f u s e r s w i t h near-optimum leng th /

d i ame te r r a t io s . In each case, t h e i n i t i a l boundary l ayer th ickness was con-

s i d e r e d small. These curves , toge ther wi th Equat ion (5 ) , can be used

e f f e c t i v e l y t o estimate t h e amount of n o i s e r e d u c t i o n f o r a g i v e n i n l e t .

C o n s i d e r t h e c a s e o f t h r e e i n l e t s w i t h a n area r a t i o of 2 .2 . Using the

upper set of curves, several p re s su re r ecove r i e s can be de t e rmined fo r

mass f low ra t ios be tween 94 percen t and 100 percent. Each point then

determines a va lue of a and consequent ly the no ise reduct ion . The approximate curve i s shown on Figure 74; it appears to estimate the noise

r e d u c t i o n q u i t e well when compared w i t h e x p e r i m e n t a l r e s u l t s .

For the case of blade passage f requency tones, Equat ion (5) should be

m o d i f i e d t o

P resen t tests and those repor ted in Reference 8 i n d i c a t e t h a t B i n c r e a s e s

wi th increas ing f requency . Fur ther work i s p r e s e n t l y b e i n g p l a n n e d t o de-

t e rmine t he i n f luence of h igh speed f low on the a t tenuat ion of v a r i o u s

f r equenc ie s .

19

8. CONCLUSION

Some g e n e r a l q u a l i t a t i v e remarks concerning high Mach number i n l e t s a p p e a r

appropr i a t e fo l lowing t he de t a i l ed examina t ion o f da t a f rom these tests as

well as f r o m o t h e r s o u r c e s ( l i s t e d i n T a b l e A of the Appendix). Whether a

choked i n l e t c a n a c t u a l l y b e u s e d d e p e n d s t o a l a r g e e x t e n t on t h e maximum

area r a t i o r e q u i r e d f o r a p a r t i c u l a r a p p l i c a t i o n . A s shown i n F i g u r e 10,

i n c r e a s i n g t h e l e n g t h o f t h e i n l e t f o r a g iven area r a t i o d o e s n o t

necessa r i ly i nc rease t he p re s su re r ecove ry nea r t r anson ic f l ow, o r a s suming

t h a t a v e r y l o n g i n l e t i s p r a c t i c a l f o r a g i v e n a p p l i c a t i o n r e q u i r i n g a

l a r g e area r a t i o f o r choked f low does not assure a h igh pressure recovery .

Thus without boundary layer control , choked inlets are a r e a - r a t i o l i m i t e d .

It a p p e a r s t h a t f o r low area r a t i o s , h i g h Mach number i n l e t s c a n b e u s e d t o

r e d u c e n o i s e w i t h s u f f i c i e n t l y h i g h p r e s s u r e r e c o v e r y a n d a c c e p t a b l e

d i s t o r t i o n i f t h e f a v o r a b l e e f f e c t s of forward speed are added which tend

t o r e d u c e t h e d i s t o r t i o n and i n c r e a s e t h e p r e s s u r e r e c o v e r y . 1

Boundary l aye r con t ro l can a lways be u sed t o i nc rease r ecove ry and r educe

d i s t o r t i o n . F o r i n l e t s w i t h h i g h area r a t i o s t h i s may be necessary. Vortex

g e n e r a t o r s are only nominal ly e f fec t ive when proper ly p laced and when t h e

degree o f separa t ion is small. I n j e c t i o n c a n b e e f f e c t i v e and is e a s i l y

added because o f the h igh pressure source ava i lab le on t h e a i r c r a f t . Suct ion i s a n a l t e r n a t i v e s i n c e t h e h i g h p r e s s u r e s o u r c e c a n b e u s e d t o

d r i v e a n e j e c t o r .

1 2

For low area ra t io s ( l e s s t han 1 .5 ) f i xed geomet ry i n l e t s w i thou t cen te rbody

appear feas ib le . This type o f in le t , however , w i l l r e q u i r e a change i n t h e

e n g i n e o p e r a t i n g c y c l e t o m a i n t a i n h i g h mass f low dur ing landing . This can

possibly be done by using a va r i ab le geomet ry ex i t nozz le wh ich can a l so be

used t o m a i n t a i n choked flow during landing.

P r i v a t e communication from B. Miller, L e w i s Research Center.

20

For higher area r a t i o s , t h e c e n t e r b o d y t y p e i n l e t a p p e a r s s u p e r i o r when

compared t o t h e c o l l a p s i n g cowl type for the same area r a t i o and l e n g t h /

d i a m e t e r r a t i o . The in le t s wi thout cen terbodies have per formed poor ly bo th

acous t ica l ly and aerodynamica l ly when compared t o t h e c e n t e r b o d y t y p e . The

t r a n s l a t i n g c e n t e r b o d y c a n a l s o b e m o d u l a t e d t o m a i n t a i n choked flow during

l and ing 13 . F o r h y b r i d i n l e t s t h e c e n t e r b o d y t y p e i n l e t p r o v i d e s a n

addi t iona l advantage because o f the curva ture . For these h igh Mach number

i n l e t s w i t h l i n e r s , t h e c e n t e r b o d y t y p e p r o v i d e s a reduced channel height

a n d a d d i t i o n a l s u r f a c e f o r t r e a t m e n t ; b o t h of t h e s e are f a v o r a b l e f a c t o r s

i n n o i s e c o n t r o l , i m p r o v i n g t h e e f f e c t i v e n e s s o f a c o u s t i c l i n e r s .

A t h igh Mach numbers , noise reduct ion appears to be f requency-dependent .

The r e l a t ionsh ip be tween t he amount of a t t enua t ion , f r equency , and flow

cond i t ions is b e i n g f u r t h e r i n v e s t i g a t e d . The p r e s e n t t r e n d s u g g e s t s t h a t

h igh Mach numbers are more e f f ec t ive i n r educ ing h igh - f r equency no i se .

Most of t he no i se r educ t ion f rom h igh Mach number i n l e t s t a k e s p l a c e a t a

value of i/ha between 90 and 100 percent. On a one-d imens iona l bas i s , th i s

means a t h r o a t Mach number v a r i a t i o n of 0.68 t o 1.0. Despi te the problem of

accuracy, the measured percent mass f l o w p l o t t e d a g a i n s t a t t e n u a t i o n i s a

be t te r parameter for compar ison be tween in le t s t h a n t h e t h r o a t Mach number.

Thus t h e e q u a t i o n r e l a t i n g n o i s e r e d u c t i o n t o Mach number

AdB = -1Olog( l / l -M)

s h o u l d b e r e w r i t t e n i n t h e f o m

where 0 < a < 1 and cx depends on the p re s su re r ecove ry .

The amount of no ise reduct ion depends on the f low downstream of t h e t h r o a t .

Thus i n l e t s w i t h h i g h l o s s e s w i l l no t p roduce l a rge no i se a t t enua t ion even

when aerodynamically choked.

21

9. REFERENCES

1 H.W. Emmons, "The T h e o r e t i c a l Flow of a F r i c t i o n l e s s , A d i a b a t i c , P e r f e c t

Gas I n s i d e a Two-Dimensional Hyperbolic Kozzle," NACA TN 1005, 1944.

2 E. Lumsdaine and L.R. Clark, "Noise Suppression with Sonic and Wear-

Son ic In l e t s , " P roceed ings of t h e Second I n t e r a g e r c y Symposiuni on

Univers i ty Research on Transportat ion Koise, North Carol ina State

University, June 5-7, 1974, pp. 432-447.

3 J.M. Wilson and C.T. S w e l l , " I n l e t Flowpa.th Design to At tenuate Super -

sonic Rotor Noise ," Genera l E lec t r ic Repcr t No. R73-AEG-412.

4 Transonic Computer Program - NASA Langley, Hampton, Vi rg in ia (used wi th

permission of E.M. Boxer).

5 E. Lumsdajme, "Resul t s of t h e Development of a Chcke.d I n l e t , " I n t e r -

Noise Proceedings, 1972, pp. 501-506.

6 C.L. Morfey ar,d F.J. F i s h e r , "Shock-Wave Radiat jon f rom a Supe.rsor?ic

Duct Rotor ," Journal of the Royal Aeronaut ical Society, Vol . 74, J u l y

1970, p. 579.

7 D.C . Matthews and R.T. Nagel, "Inlet Geometry and Axial Mach Number

E f f e c t s on Fan Poise Propagat icn ," AIM Paper No. 73-1022, AIM. Aero-

Acoust ics Cocference, Seat t le , Washington, October 1973.

8 Y. Tuan, "A Survey of Sonic In le t Concepts for In le t Noise Reduct ion ,"

Boeing Document D6-40573, May 1972.

22

9 Anon., P ro jec t S t a tus Repor t PE-8217-R6, Airesearch Manufacturing

Company of Arizona, Garrett Corpora t ion , November 1971.

10 M.J.T. Smith and M.E. House, " I n t e r n a l l y Genera.ted Noise from Gas

Turbine Engines - Measurements and Predictions," ASME Paper No. 66-GT/

N-43, Yarch 1966.

11 E. Lumsdaine, "Fan Noise Reduction at Subsonic and Supersonic Tip

Speeds with High Mach Number In le t s , " In te r -Noise Proceedings , Sendai ,

Japan , 1975.

12 E . Lumsdaine ar,d A . Fathy, "Theoretical Study of Boundary Layer Control

by Elowing f o r Axisymmetric Flow and I t s Appl ica t i .on to the Sonic

Ir,let," Eng inee r ing Expe r imtx t S t a t ion Bu l l e t in 1 6 , South Dakota State

Universi ty? October 1970.

1 3 J. Jibben and E . Lumsdaine, "An Automatic Control System for Sonic

I n l e t s , " t o b e p r e s e n t e d a t t h e A S E Autcnat ic Contro1.s Conference,

Houston, Texas, December 1.975.

14 David Chestnutt and Lorenzo R. Clark, "Noise Reduction b57 Means of

Variable-Geometry I n l e t G u i d e Var?es i n a Cascade Appzratus," NASA Tp.I

X-2392 October 1971.

15 F. Klujber , K.C. Eosch, R.W. Demetrick, W.R. Robb, "1nvestigati .on of

No i se Suppres s ion by S0~ j . c In l e t s fo r Turbofan Eng ines - Final Report ,

Vol.umes I and 11," Eoeing Document D6-40855, NASA CR-1211.26, CR-121127,

Ju ly 1973.

23

I FLOW CONFIGURATIONS ~~ -

FIG. 1 WALL AND CENTERLINE MACH NUMBERS FOR POTENTIAL FLOW IN TWO DIMENSIONS

24

0

- 10

-20

-30

-40

-50

- AdB = 10 log (-) 1 l-Mn

" "-

\ I

0 LID = 0.67

0 LID = 1.76

LID = 3.15

0 NASA- Ches tnu t t & Clark (Throat 4 Mach Number)

-" Empir ica l Equat ion

I , 0.8 0 . 9 1.0 M~~~ AVERAGE MAXI" AXIAL MACH NUMBER

1 .2

FIG.2 INFLUENCE OF LENGTH/DIAMETER RATIO ON NOISE ATTENUATION (EJECTOR SOURCE)

4 4

5 \

3.5

3.0

2.5

CONFIGURATION NO. 1

25.4 cm (10 in.) -

- 20.3 cm ( 8 in.)

12.7 cm ( 5 in.)

2.0

1 . 5

1.0

/ / I / Centerbody Fully Retracted

0 20 30 40 50 60 7 0 80 DISTANCE, CM

.c .

FIG.3 SCHEMATIC ANTI AREA DISTRIBUTION OF TRANSLATING CENTERBODY INLET CONFIGURATION 1

26

3.0

2.5

0

3 2.0

4

1.0

CONFIGURATION NO. 2

7 . 6 cm -

0.5 -6.35 0 10 20 30 40 50

I I I I I I ,

DISTANCE, CM

FIG-4 SCHEMATIC AND AREA DISTRIBUTION OF TRANSLATING CENTERBODY INLET CONFIGURATION 2

(4.5 in.)

(3.0 in.)

(1.5 in.)

27

0 ! 'tl I n 1 I4

t

CONFIGURATION NO. 3

3.0

2 . 5

0 H

3 2.0

1.0

0 . 5 . - 2

i

( 4 . 5 in.)

( 3 . 0 in.)

(1.5 in.)

I . o 10 20 30 40 50

1 1 1 1 I )

DISTANCE, CM

FIG.5 SCHEMATIC Ah3 AREA DISTRIBUTION OF TRANSLATING CENTERBODY INLET CONFIGURATION 3

28

...

"

CONFIGURATION NO. 4

0 U

a! a!

5 \

/ T

11.4 cm (4.5 in.)

2 . 0 -

I / / 0 cm

0.5 I I 1 1 I I -

O 10 20 30 40 50

DISTANCE, CM

FIG.6 SCHEMATIC AND AREA DISTRIBUTION OF TRANSLATING CENTERBODY INLET CONFIGURATION 4

29

50.8 cm-J ,&

Microphones

b - h LAY

3.0

2.5

0 H

3 2.0

c 1 4 1.5

1.0

0.5

CONFIGURATION NO. 5

1 1 I I"-I - 10 20 30 40 50-

DISTANCE, CM

FIG.7 SCHEMATIC AND AREA DISTRIBUTION OF FIXED GEOMETRY (COLLAPSING COWL) INLET CONFIGURATION 5

30

t- 50.8 cm -4

Microphones

3 7 1

CONFIGURATION NO. 6

2 . 5

2 . 0

0 H

3 -4 3 1 . 5

5 \ -4 -4

1.0

0.5 I I I I 1 - -

31

CONFIGURATIONS NO. 7,8,9

2 .5

2.0

0 H

3 2 3 1.5 5 \ -4 4

1.0

0.5

- 6 . 10 20 30

DISTANCE, CM 40 50

32

100

95

90

85

80

75

70

-

r -10 AdB

-20 AdB

-30 AdB

-40 AdB

100

95

I I I I L I I I

2 4 6 8 - -

LENGTH/FLOW DIAMETER

I I I I L t

- 1 1 2 3 4

LENGTHIDIANTER RATIO, L/D

M

0.

0.

0.

0.

1.

I I 1 I I

2 4 6 L+

8 LENGTH/FLOW DIAMETER

I

I I I I t 3 1 2 3 4

LENGTH/DIAMETER RATIO, L/D

FIG.10 AERODYNAMIC AND ACOUSTIC PERFORMANCE OF INLETS WITH DIFFERENT LENGTH/DIAMETER RATIOS

w

5.0

4.0

CI (d 0 &l

4 5 3.0 \

1-I CI

1 4

2.0

4 1.0

/

6

0 Leon Sap i ro (M=O. 2)

@ Leon S a p i r o (M=0.6)

@ Leon Sapiro (M=l. 0)

6 Runs tad le r (M=0.2)

Runs tad le r (M=0.6)

d Runs tad le r (M=1.0)

a James P. Johnston

@ Reneau e t a l . (M=l. 0)

A R.V. Van Dewoestine (M=O 7)

Rhoades

6 Reneau e t a l . ( M = l . 0)

0 1 2 3 4 5 6 7 a 9 10 DIFFUSER LENGTH/EXIT FLOW DIAMETER, L/D,~~

F I G . l l OPTIMUM L/D FOR A GIVEN AREA RATIO AT DIFFERENT THROAT MACH NUMBERS FOR CIRCULAR AND SQUARE DIFFUSERS

Figure 10 and 72 In te r -Noise x Proceedings ( M t h = l .O)

0 Genera l Electric (Mth=l.O)

Boeing Document D6-40855 ' ( M t h = l . 0)

0 1 2 3 4 5 6 7 8 9 10

DIFFUSER LENGTH/EXIT FLOW DIAMETER, L/D

FIG.12 OPTIMUM L/D FOR A GIVEN AREA RATIO AT DIFFERENT THROAT MACH NUMBERS FOR ANNUJAR DIFFUSERS

0.8

0.7

T h e o r e t i c a l Mach N u m b e r D i s t r i b u t i o n

A C o w l

0.1 - I I , I I I I I I I 1 1 I I I I 1 I I , - 10 0 10 20 30 40 50 60 70 80 "

t DISTANCE FROM LEADING EDGE OF COWL, CM 1.0

0.9

0 . 8

0 . 7

E x p e r i m e n t a l Mach Number D i s t r i b u t i o n

0 . 4 1 e 0- d

A A

0.3

0.1-

- 0 .2

-

I I I I I I I I I I I I I I I I ~~ I 1 1 c

FIG.13 THEORETICAL POTENTIAL FLOW ANI) EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 0 CM DISPLACEMENT

36

-

Streamlines

Centerbody I

l l l l r , r . I I s I I I

I I

0.9 t Theoretical Mach Number Distribution

l r I 1 I I I I I c

-10 0 10 20 30 40 50 60 70 80

1.0 I 1 DISTANCE FROM LEADING EDGE OF COWL, CM

0.9 1 0 . 8

0.7

0.6

0.5

0.4

0.3

Experimental Mach Number Distribution

FIG.14 THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 5.1 CM DISPLACEPENT

37

1.0

0.9

0.8

0.7

0 .6

0.5

0 .4

0.3

0.2

0.1

t t T h e o r e t i c a l Mach Number D i s t r i b u t i o n ~~

A C o w l

.- 0 Cente rbody

I t I I 1 I I I I I I , I I I I , -10 0 10 20 30 40 50 60 70 80

DISTAXCE FROM LEADING EDGE OF COWL, CM 1.0 " 0.9 -

0.5 -

E x p e r i m e n t a l Mach Number D i s t r i b u t i o n

0.1 - 1 I I I 1 I I I I I I I I I I I I b

F I G S 1 5 THEORETICAL POTENTIAL FLOW EXPERIMENTAL RESULTS FOR CONFIGLJRATION 1 WITH CENTERBODY AT 12 .7 CM DISPLACEMENT

38

I

Streamlines

Centerbody I

I I

I I I I l l I I

Theoretical Mach Number Distribution

a Cowl

0 Centerbody

I I I I I I ,I I I I I i I I 1 I I I I I

4 DISTANCE FROM LEADING EDGE OF COWL, CM -10 0 10 20 30 40 50 60 70 80

1.2

1.0

- 1.1

-

Experimental Mach Number Distribution n.9 - 0.8

0.6

- 0.7

-

- 0.5 - 0.4 -

FIG.16 THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 17.8 CM DISPLACEMENT

39

- Streamlines

Centerbody I I

X

Y

Theoretical Mach Number Distribution

A Cowl

0 Centerbody

I I I I 1 I I I I I I I I 1 I I L -20 0 10 20 30 40 50 60 70 80

DISTANCE FROM LEADING EDGE OF COWL, CM

A

I I I I I 1 I I I I I I I I 1 I I L -20 0 10 20 30 40 50 60 70 80

DISTANCE FROM LEADING EDGE OF COWL, CM

A

0.6 - 0.5

- 0 . 4

- Experimental Mach Number Distribution

FIG.17 THEORETICAL POTENTIAL FLOW AND EXPERIMENTAL RESULTS FOR CONFIGURATION 1 WITH CENTERBODY AT 20.3 CM DISPLACEMENT

40

F I G . 1 8 NASA-LANGLEY RESEARCH CENTER ANECHOIC-CHAMBER TRANSONIC COMPRESSOR FACILITY

41

c- W

Instrumentation Compressor

Static Pressure Taps

Static Pressure

Instrumentation

spOO1 7 Static Pressure Taps

P Flow P ”

FIG.21 SCHEMATIC OF INLET CONFIGURATION 5 WITH AERODYNAMIC INSTRUMENTATION

1.50

1.40

u "

\ PI N u

PI

0 H

2 1.30

!i i cn

PI

3

3 Fr,

IJ 1.20 0 H

1.10

100 P e r c e n t of Maximum RPM

Compresso r ope ra t ing po in t s f o r d a t a a c q u i s i t i o n :

0 C o n s t a n t P r e s s u r e R a t i o

A Cons tan t RPM

X Near Surge

I I 1 I I I - - 0.4 0.5 0 . 6 0.7 0.8 0.9 1.0

AVERAGE THROAT MACH NUMBER BASED ON ONE-DIMENSIONAL FLOW

FIG.22 ILLUSTRATION OF INLET AND COMPRESSOR OPERATING CONDITIONS FOR DATA ACQUISITION

F9 a a

rl

3 0

w

I Boom M i c r o p h o n e P o s i t i o n a t 0 deg

+lot 6K 8K 10K 1 2 K

0 I I I I

-10

-20 X Centerbody a t 1 2 . 7 cm

A Centerbody a t 17.8 cm

Centerbody a t 20.3 cm

0 C e n t e r b o d y a t 2 5 . 4 cm

- 30

-40

FIG.23 NOISE ATTENUATION FOR INLET CONFIGURATION 1

12

10

U w \ rn 0 8 M

t2 r n 6

3 E i i 4 0 U

*E

2

0

0 Centerbody a t 0 cm

x Centerbody a t 12.7 cm


a Centerbody a t 20.3 cm

0 Centerbody a t 25.4 cm

4K 6K 8K 10K 12K 14K 16K 18K 20K 22K 24K

CORRECTED RPM

FIG.24 MASS FLOW FOR INLET CONFIGURATION 1

1.4 t 1.2 -

8 z

5 1.0 3

-

2 z

!i O a 8 - c3 W

2 0.6 - z u >

0.4 -

0.2 -





@ Centerbody a t 25.4 cm

0 1 4K 6K 8K 10K 12K 14K 16K 18K 20K 22K 24K 26K

CORRECTED RPM

FIG.25 AVERAGE MAXIMUM MACH NUMBER OF INLET CONFIGURATION 1

c U3

J

100

95

90

85

80

75


X Centerbody a t 1 2 . 7 cm

A Centerbody a t 1 7 . 8 cm


0 Centerbody a t 2 5 . 4 cm

c 20 dB

.y$ \j, ~ 8 . 4 dB 23 dB

CORRECTED RPM

FIG.26 PRESSURE RECOVERY FOR INLET CONFIGURATION 1




Centerbody a t 20.3 CIL

0 Centerbody a t 25.4 cm

" 4K 6K 8K 10K 12K 14K 16K 18K 20K 2 2K 24K 26K

CORRECTED RPM

2 -

1 -

0 I *

4K 6K 8K 10K 12K 14K 16K 18K 20K 2 2K 24K 26K

CORRECTED RPM

FIG.27 PRESSURE DISTORTION FOR INLET CONFIGURATION 1

m a a

E 0 rn

+10 't Boom Microphone Posi t ion a t 0 deg

-10 -

.-20 -

-30 -

-40 - I

@ Centerbody at 0 cm

X Centerbody at 3.8 cm

A Centerbody a t 7 . 6 cm

Centerbody at 1 1 . 4 cm


0 Centerbody at 0 cm

X Centerbody at 3.8 cm

A Centerbody at 7.6 cm

Centerbodv at 11.4 cm

2233.5 & 38.5 dB

dB

4K 6K 8K 10K 12K 14K 16K 18K 2 OK 2 2K 24K -

CORRECTED RPM


I X Centerbody a t 3.8 cm

1.0 - I

0.8 -

0.6 -

0 . 4 -

0.2 -

0' I 1 I I I I I I I I L

4K 6K 8 K 1 OK 12K 14K 16K 18K 2 OK 2 2K 24K

CORRECTED RPM

FIG.30 AVERAGE MAXIMUM MACH NUMBER OF INLET CONFIGURATION 2

1

100

98

96

64

92

90

88

86


X Centerbody a t 3 . 8 cm


Centerbody a t 11 .4 cm

4 K 6K 8K 1 OK 12K 1 4 K 16K 18K 20K 2 2K 2 4 K

CORRECTED RPM


H n

@ Centerbody a t 0 cm



1 2

10 .

8 ,

6 '

4

2

a Centerbody a t 11.4 cm

P

0 1 I I I I I I I I I I I F

4K 6K 8 K 10K 12K 14K 16K 18K 20K 2 2K 2 4K 2 6K

CORRECTED R P M


I

+10

0

- 10

-20

-30

-40

Boom M i c r o p h o n e P o s i t i o n a t 0 deg

I

0 C e n t e r b o d y a t 0 cm


A C e n t e r b o d y a t 7 . 6 cm

Centerbody a t 1 1 . 4 cm


12

1 0

8

6

4

2

0


x Centerbody a t 3 .8 cm

A Centerbody a t 7 .6 cm

Centerbody a t 11.4 cm 30 dB

4K 6K 8K 10K 12K 14K 16K 1 8 K 20K 22K 24K CORRECTED R P M


1 . 4

1 . 2

1 1.0 E 3 0.8

$

3 ’ 0.6 w

W ’ 0 . 4 c1

0 . 2

0


X Centerbody a t 3.8 cm



6K

FIG. 35

8 K 1 OK 12K 14K 16K 1 8 K 20K

CORRECTED RPM

AVERAGE MAXIMUM MACH NUMBER FOR INLET CONFIGURATION 3

2 2K

100

98

96

94

92

90

88

86

0 Centerbod .y a t 0 cm



Centerbody a t 1 1 . 4 cm

4K 6K 8K 10K 12K 14K 16K 18K 20K 2 2K 24K

CORRECTED RPM


1 4

1 2

10

8

6

4

2

0

Centerbody a t 0 c m

Centerbody a t 3 . 8 c m


Centerbody a t 1 1 . 4 c m

I I I t r

4K 6K 8 K 1 OK 12K 14K 16K 1 8 K 2 OK 2 2K 2 4K 2 6K

CORRECTED RPM


a E4

Q +IO 1 Boom Microphone Position a t 0 deg

0

- 10

-20

-30

-40


x Centerbody at 3.8 cm



12

10

8

6

4

2

C







1 I I 1 I I I I I I ,

4K 6K 8K 1 OK 12K 1 4 K 16K 1 8 K 2 OK 2 2K 24K

CORRECTED RPM


I @ Centerbody a t 0 cm

X Centerbody a t 3 .8 cm


I I I 1 I I I I I 1 c 4K 6K 8K 10K 12K 14K 16K 1 8 K 20K 22K 24K

CORRECTED RPM

FIG. 40 AVERAGE MAXIMUM MACH NUMBER FOR INLET CONFIGURATION 4

I

100

98

96

94

92

90

88

86

Centerh lody

Centerbody a t 1 1 . 4 cm \ I 1 3 7 dB

I I I I I I I I I I m 4K 6K 8K 1 OK 12K 14K 16K 1 8 K 20K 2 2K 24K

CORRECTED RPM


14

12

10

8

6

4

2

0

Centerbodv a t 0 cm

Centerbody a t 3 . 8 cm


I I I I I I I I I I I .

4K 6K 8K 10K 12K 14K 16K 1 8 K 2 OK 2 2K 2 4K 26K-

CORRECTED RF”


+10

I

0 -

- 10

-20

-30

- 40





FIG.43 EFFECT OF PERCENTAGE OF MAXIMUM MASS FLOW ON NOISE ATTENUATION FOR INLET CONFIGURATION 2

L

1 I

+10 -

m E 4 rn

0

-10

-20

-30

-40




0 Centerbody a t 1 1 . 4 cm

FIGS44 EFFECT OF PERCENTAGE OF MAXIMUM MASS FLOW ON NOISE ATTENUATION FOR INLET CONFIGURATION 3

+lot

a m a .. IJ

3 w w I-1

F I G . 45



A Centerbody at 7.6 cm

& A 0 30 40 50 60 7 0 80 4 90 100

I --

As

EFFECT OF PERCENTAGE OF MAXIMUM MASS FLOW ON NOISE ATTENUATION FOR INLET CONFIGURATION 4

i max - i , %

.. .

Boom Microphone Position at 0 deg "10

--I 0 rr

w A U ,

n 20K 2 2K 24K RPM 0 I I I I I I I C

6K 8K 10K 12K

-10 -

- 2 0 -

-30 -

-40 - 'I

18K 20K 2 2K 24K RPM I I I I I C

0 Configuration 5

X Configuration 6

FIG.46 NOISE ATTENUATION FOR INLET CONFIGURATIONS 5 AND 6

0 C o n f i g u r a t i o n 5

X C o n f i g u r a t i o n 6

4.0 dB

11.5 dB

CORRECTED RPM

FIG.47 MASS FLOW FOR INLET CONFIGURATIONS 5 AND 6

1.4

1.2

1.0

0.8

0.6

0.4

0.2

0

Q Conf igura t ion 5

X Conf igura t ion 6

4K 6K 8K 1 OK 12K 14K 16K 18K 20K 2 2K 24K

CORRECTED RPM

FIG.48 PEAK WALL MACH NUMBER OF INLET CONFIGURATIONS 5 AND 6

FIG.49 PRESSURE RECOVERY FOR INLET CONFIGURATIONS 5 AND 6

25

20

15

10

5

0



4K 6K 8K 10K 12K 14K 16K 18K 2 OK 2 2K 24K 26K

CORRECTED RPM

FIG.50 PRESSURE DISTORTION FOR INLET CONFIGURATIONS 5 AND 6

+10

0 0 x n

i i i I I I 30 40 50 60 70 80

Q Configuration 5

X Configuration 6

lil

max - fi , %

- 30 0

I? FIG.51 EFFECT OF PERCENTAGE OF MAXIMUM MASS FLOW ON NOISE ATTENUATION FOR INLET CONFIGURATIONS 5 AND 6


x C o n f i g u r a t i o n 8

A C o n f i g u r a t i o n 9

0 I I I I I I I I I I I >

4 K 6K 8K 10K 1 2 K 1 4 K 1 6 K 18K 20K 2 2K 2 4 K 2 6 K

CORRECTED RPM

FIG.53 PEAK WALL MACH NUMBER OF INLET CONFIGURATIONS 7, 8 AND 9

100

98

96

94

92

90

88

86

@ C o n f i g u r a t i o n 7


C o n f i g u r a t i o n 9

4K 6K 8K 10K 1 2 K 14K 16K 1 8 K 20K 2 2K 2 4K 2 6K

CORRECTED R P M

FIG.54 PRESSURE RECOVERY FOR INLET CONFIGURATIONS 7 , 8 AND 9

0 Configuration 7

x Configuration 8

A Configuration 9

4 K 6 K 8 K 1 0 K 1 2 K 1 4 K 1 6 K 1 8 K 2 0 K 2 2K 24K 26K

CORRECTED RPM

FIG.55 PRESSURE DISTORTION FOR CONFIGURATIONS 7 , 8 , 9

I I - "

110

100

2 a cl w > W cl

I

g 90 H 0 z I 4

3 W iz

80

70

Boom Microphone Position at 45 deg

X Configuration 8

0 Configuration 1 (Centerbody at 0 cm)

-. I I I I 1 c - 5,000 10,000 15,000 20,000 25,000

R P M

FIG.56 COMPARISON OF OVERALL NOISE ATTENUATION CHARACTERISTICS OF INLETS OPERATING AT SUBSONIC MACH NUMBERS

79

11

10

9

e(

7(

Boom Microphone Position at 45 deg

F I G . 5 7 COMPARISON OF BLADE PASSAGE FREQUENCY ATTENUATION CHARACTERISTICS OF INLETS OPERATING AT SUBSONIC MACH NUMBERS

80

I

Graph Centerbody Posi t ion - RPM

A dB! Boom Microphone Posi t ion a t 45 deg No Centerbody 17753

I

5c

4c

30

20

10

0

"" 0

""_ 3.8 cm

17226

17492

17400 I 1

- . . ,_ 11.4 cm 17520 -. 1.- c I1 "'1 _._

.... -L ...

\

I I

5 20 50 200 500 2K 5K 20 50 K FREQUENCY, 'rIERTZ

FIG.58 FREQUENCY SPECTRA FOR CONFIGURATION 2 AT 17,500 RPM

A dB

60

50

40

30

03 N

20

10

d BA Without Centerbody, 20254 RPM, Unchoked

I-

L Cente rbody Pos i t i on 0 cm, 20112 RPM, Choked 7

- 1

- -

4

I i

I I I I I 1 I I

5 20 50 200 5 00 2 K 5 K 2OK 50 K

FREQUENCY, HERTZ

FIG. 59 FREQUENCY SPECTRA FOR CONFIGURATION 2 AT 20,100 R P M

Q, W 0 Short Bellmouth

x Simulated Landing Lip

A F l i g h t L i p (Take-Off)

“4 Centerbody Trans la t ion : 0 cm

FIG.60 INFLUENCE OF L I P GEOMETRY ON NOISE ATTENUATION

100'

95

z B

u w

90 .,

80

75

0 Short Bel lmouth

X Simulated Landing

A Fl ight L ip (Take-Off )

Cen te rbody T rans l a t ion : 0 cm

+5 0 -5 -10 -15 -20 -25 -30 -35 -40 CHANGE I N OVERALL SOUND PRESSURE LEVEL, AdB

FIG.61 INFLUENCE OF LIP GEOMETRY ON NOISE ATTENUATION AND PRESSURE RECOVERY

H z w U p:

90 0 Short Bellmouth

X Simulated Landing

F l i g h t L i p (Take-Off)

Centerbody Trans la t ion : 3.8 cm

100 - 0

95 -

-

-

75 I I I I I I I I I C

+5 0 -5 -10 -15 -20 - 25 -30 -35 -40 CHANGE I N OVERALL SOUND PRESSURE LEVEL, AdB

80


0 Short Bel lmouth

x Simulated Landing

A Fl ight L ip (Take-Off )

Cen te rbody T rans l a t ion : 7 . 6 cm

-A

0 -5 -10 - 15 -20 -25 -30 -35 -40 ~~

7

CHANGE I N OVERALL SOUND PRESSURE LEVEL, AdB

FIG.63 INFXUENCE OF L I P GEOMETRY ON NOISE ATTENUATION AND PRESSURE RECOVERY

+5

0 Short Bellmouth

x Simulated Landing

Centerbody Trans la t ion : 11 .4 cm

0 -5 - 10 -15 - 20 -25 -30 -35 CHANGE I N OVERALL SOUm.PRESSURE LEVEL, AdB


111

lo (

9(

8C

70

110

100

90

80

70

0 Inlet Without Centerbody

Configuration 2 1 I I I I

Boom Microphone Position at 0 deg 0 Inlet Without Centerbody

X Centerbody at 0 cm

P

Configuration 4 x RPM

5,000 10,000 15,000 20,000 25,000

FIG.65 INFLUENCE OF CENTERBODY ON NOISE LEVEL

88

90" 80" 70" 60" 50" 40 "

30"

0 5,000

X 10,000 RPM

15,000 RPM

0 20,000 R P M

0 25,000 RPM

0"

30"

20"

O0

30

20 "

10 "

O 0

F I G . 6 6 POLAR GRAPHS SHOWING THE INFLUENCE OF LIP GEOMETRY OF INLETS OPERATING AT SUBSONIC MACH NUMBER

89

1.4

s 1.3

3

i E! Cn

PI 1.2

3 cr

1.1

C o n f i g u r a t i o n 1 86

Centerbody T r a n s l a t i o n : 20.3 cm 90

96 I

84 dBA

< \ \ \ \

x 24,000 R P M

0 22,500 RPM

21,000 RPM

A 20,000 RPM

15,000 RPM

1.0- L I I I I 1 I

0 2 4 6 8 10 1 2 b-

r i ~ CORRECTED MASS FLOW RATE, KG/SEC

FIG.67 INFLUENCE OF COMPRESSOR PRESSURE RATIO ON NOISE

PWI

T 5dI

I -

~

Rth/Rfan'0.89

(L/D=l. 14)

C o n f i g u r a t i o n 1 0 ( P r e s e n t T e s t s )

kh=0.89 (L/D=l. 14)

A General E l e c t r i c (No.2)

General E l e c t r i c (No.3)

0 General E l e c t r i c (No.5)

Rth = T h r o a t R a d i u s of I n l e t

Rf an = R a d i u s a t F a n F a c e

0 I I

1.1 1.2 1.3 M' t fP

0.56 0.605 0.655 Mroot ROTOR RELATIVE MACH NUMBER

FIG.68 COMPARISON OF NOISE ATTENUATION CHARACTERISTICS OF INLETS OPERATING AT HIGH SUBSONIC MACH NUMBER WITH COMPRESSOR OPERATING AT SUPERSONIC ROTOR TIP SPEED

91

Present Tests: A2 /Al=l. 4 L/D=1.68

A2/Al=1.8 LlDO1.68

L/D=l. 68 I3 A2IAle2.2

x A21A1-2.6 L/D=l. 68

F4 0

1 N

PI 0

W

0

r(

\ N 0

PI

F4

100

95

90

Contoured Lip

85 +10 0 -10 -20 -30 -40

CHANGE IN OVERALL SOUND PRESSURE LEVEL, AdB

85 I I I I I I

+10 0 - 10 -20 -30 -40 CHANGE IN OVERALL SOUND PRESSURE LEVEL, AdB

Boeing Tests: A2/A1=1.18 L/D=I.o

v A2/A1=1.64 L/D=l. 0

8 A2 /Al=l. 64 LfD=l. 3

A2 /AI =1.4 ' L/D=1.68 A A~/A1=1.8

L/D=1.68 A2/A1=2.2 L/D=1.68

FIG.69 EFFECT OF AREA RATIO ON NOISE ATTENUATION AND PRESSURE RECOVERY

92

Po /Po PRESSURE RECOVERY, PERCENT

FIG.70 INFLUENCE OF PRESSURE RECOVERY ON THE EXPONENTIAL PARAMETER a

1.0

0.5

0.4

c1

0.3

0.2

0.1

0

93

-30

w I

-40

@ David Chestnut , Lorenzo R. C l a r k , NASA TM X-2392, Oct. 1971

S t a t i o n a r y Uncambered I n l e t Guide Vanes

-10

-20

-30

-40

FIG.71 COMPARISON BETWEEN EMPIRICAL EQUATIONS AND EJECTOR TEST DATA

I I I I I I T

----"""- ""- "

10 l og [ 1 1 - '"/mma, . . )l/a 1

1 1 - M 10 log["]

0 c

0 F. Klu jber , J .C. Bosch, Boeing Company, NASA CR-121126, July 1973

Trans la t ing Centerbody, LID = 1.0, Run No.6, Model 4

-10

-20

-30

-40

FIG.72 COMPARISON BETWEEN EMPIRICAL EQUATIONS AND COMPRESSOR TEST DATA

Dumped d i f f u s e r

3 I I I .

2.0 3.0 4.0 AREA RATIO

F I G . 7 3 P R E S S U R E RECOVERY FOR DIFFUSERS WITH DIFFERENT AREA RATIOS

96

F9 s

-10 w

d

w 30 -30

w w

!! U

-40

@ Exper imenta l Data (A2/A1 = 2.2)

-30

-40

FIG.74 COMPARISON OF EMPIRICAL EQUATIONS WITH EXPERIMENTAL RESULTS

A P P E N D I X

TABLE A - REVIEW OF WORK ON SONIC AND NEAR-SONIC INLETS

Year Author (s) Paper Type of Test

1961 Sobel and Well iver CURTISS-WRIGHT

~~ ~~~ ~ ~ ~

S o n i c i n l e t tests with centerbody on compressor and Olympus-6 t u r b o j e t

Noise-Control, Vo1.7, No.5

1964 McKaig, BOEING Document T6-3173 SST-type i n l e t on 5-75 engine

1966 Sawhill , BOEING Document D6A 10155-1 E j e c t o r test wi th 12.7-cm SST-type i n l e t

E j e c t o r tests wi th 12.7-cm SST-type i n l e t 1966 Anderson, BOEING Document D6A 10378-1 TN

1967 Cawthorn e t a l . , NASA TN D-3929 SST-type in l e t w i th V ipe r -8 j e t engine

1968 E.B. Smith et a l . GENERAL ELECTRIC

TR DS-68-7, Con t rac t FA65WA-1236, FAA

Model cascade

1968 Chestnut t NASA

TN D-4682 E j e c t o r tests with cambered and uncambered a i r - f o i l s

1968 Higgins et a l . , BOEING SP-189, pp. 197-215, NASA

Document D6-23469

JT3D engine wi th cont rac t ing cowl

1969 J . N . Smith and Higgins BOEING

JT3D engine

1970 Putnam and R.H. Smith NASA

TN D-5692 S t a t i c test wi th XB-70 a i r p l a n e

1971 Chestnut t and Clark NASA

TM X-2392 Var i ab le geomet ry ca scade i n l e t tests w i t h e j e c t o r

1972 Anderson et a l . , BOEING G r i d a n d r a d i a l t y p e v a n e i n l e t w i t h f a n Document D6-40208

72 Inter-Noise Proceedings 1972 Lumsdaine SOUTH DAKOTA STATE U., UNIVERSITY OF TENNESSEE

1973 Klujber , BOEING

Model tests o f s e v e r a l i n l e t s w i t h e j e c t o r

Transonic 0.35-m d i a . € a n u s i n g i n l e t s with low area r a t i o (max. A2/A1 = 1.6)

Document D6-40855 (Vols. I, 11, 111)

(cont . )

TABLE A con t . )

Year Author ( s ) Paper Type of Test

1973 Miller and Abbott NASA

TM X-2773 S m a l l f a n w i t h c r o s s f l o w a t v a r i o u s a n g l e s w i t h t r a n s l a t i n g c e n t e r b o d y ( two pos i t i ons on ly )

Tests wi th Lang ley t r anson ic compresso r w i th , expand ing cen te rbody (A /A = 2.6)

S tudy of engine var iab le geometry sys tems w i t h h i g h Mach number i n l e t s - c o l l a p s i n g cowl i n l e t recommended

2 1

197 3 Kutney GENERAL ELECTRIC

Work i n p r o g r e s s ( a l s o G.E. Repor t No. R-73-AEG-412)

1973 Compagnon GENERAL ELECTRIC

NASA CR-134495

1974 Lumsdaine UNIVERSITY OF TENNESSEE

Second Interagency Symposium on Un ive r s i ty Resea rch i n T r a n s p o r t a t i o n N o i s e

Tests wi th Langley t ransonic compressor w i t h v a r i o u s area r a t i o s a n d v a r i o u s t y p e s of i n l e t s ( m a x . A2/A1 = 3.6)

F i r s t test o f a n a u t o m a t i c c o n t r o l s y s t e m d e s i g n e d f o r t h e c h o k e d i n l e t

P 0 0

1974 Lumsdaine and Jibben SOUTH DAKOTA STATE U . , UNIVERSITY OF TENNESSEE

To b e p r e s e n t e d at the 1975 ASME Automatic Control Conf . , Houston, Texas

AIAA paper 74-91 1974 Groth NASA

T r a n s l a t i n g c e n t e r b o d y w i t h r a d i a l v a n e s t e s t e d w i t h a 5-85 eng ine

1974 Koch e t a l . ALLISON D I V I S I O N , GM

AIAA paper 74-1098 F ixed geomet ry i n l e t t e s t ed w i th mode l f an (115 scale of advanced fan)

1974 Lumsdaine and Clark UNIVERSITY OF TENNESSEE and NASA

Second Interagency Symposium on Un ive r s i ty Resea rch i n T r a n s p o r t a t i o n N o i s e

R e s u l t s o f c h o k e d i n l e t s t e s t e d w i t h 30.5-cm f a n

F 0 1975 Y

S o n i c a n d n e a r - s o n i c i n l e t tests w i t h NASA Langley t ransonic compressor

E f f e c t o f l i p d e s i g n o n a c o u s t i c a n d a e r o - dynamic performance of high Mach number i n l e t s , tests a t d i f f e r e n t a n g l e s of a t t a c k i n w i n d t u n n e l , s i r e n s o u r c e

Lumsdaine UNIVERSITY OF TENNESSEE

75 In te r -Noise Proceedings

B. Miller e t a l . NASA

NASA TM X-3222

with mach number inlets - ntrs.nasa.gov · high mach number inlets edwmd lunzsdaine, jenn g....

Documents