NACA
RESEARCH MEMO
AN 8-FOOT AXBY&%BWCAL FIXED NOZZLE FOR SUBSONIC MACH
1
.
NUMBERS
., ,y-, -'-
*d.
J..
UP
?r, 0.99
AND FOR A SUPERSOW
L
- .:
,m, .L&J&p.-r+
- .
..
,
-a
~,<G@iie&
A-J~-?<!
, , MACH
? rc2c--:-;,L=z
NUMBER
- , # , , - OF
-7;. -.I:, .
, 7-.--.-'TALx:-;
:~-T;>C;:$:~:I‘7,--..
z-
By ~ i r & l -~~i t.c-h i eRay
,
H. ~ r i k ' t , '
and Marshall P. Tulin
Langley Aeronautical Laboratory
I a q l e y A i r Force Base, Va.
- .;-,<:$I-,--,+
7z+h- 1:. y e --,+: .+,-I.K-- > v<Ly
.=-.:?-c-?i;
-
.-
i_L---
NATIONAL ADVISORY ~
~
~
~
i
l
NACA RM L50A03a
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS
RESEARCH MEMORANDUM
AN 8-FOOT MISYMMETRICAL FIXED NOZZLE FOR SUBSONIC MACH
NUMBERS W TO 0.99 AND FOR A SUPERSONIC
MACH NUMBER OF 1.2
By V i r g i l S. Ritchie, Ray H. Wright,
and Marshall P. Tulin
SUMMARY
The design and operation of a f i x e d nozzle of a x i a l symmetry f o r
high-subsonic Mach numbers and f o r a supersonic Mach number of 1.2 was
investigated i n connection with t h e conversion of a l a r g e high-speed
subsonic wind tunnel t o transonic operation. The nozzle was c a r e f u l l y
designed by p o t e n t i a l - f low theory and was adjusted f o r boundary-layer
development. The r e s u l t s of flow surveys i n t h e nozzle indicated t h a t
the uniformity of t h e flow i n t h e supersonic t e s t section was s u f f i c i e n t f o r model t e s t i n g . I n the subsonic t e s t s e c t i o n provided i n t h e
region of t h e nozzle t h r o a t , t h e flow was of remarkable uniformity and
permitted t h e t e s t i n g of small models at Mach numbers a s great a s 0 .gg.
q
.
Small flow i r r e g u l a r i t i e s , which were propagated along Mach l i n e s ,
tended t o become concentrated near the nozzle a x i s . Reduction of some
waviness i n the surface of t h e nozzle by amounts of t h e order of
0.006 inch r e s u l t e d i n improvement of t h e flow, and deviations from t h e
design Mach number a t t h e nozzle a x i s d i d not then exceed 0.02. Small
surface i r r e g u l a r i t i e s of the nature of roughness, cracks i n the surface
of the p l a s t e r l i n e r , and small d i s c o n t i n u i t i e s i n slope produced no
noticeable e f f e c t on the flow and were presumed masked by t h e t h i c k
boundary l a y e r .
I
-
Humidity e f f e c t s were controlled by heating of the flow mixture.
No condensation shocks could be detected even with considerable cooling.
The slow formation of fog, which occurred with supercooling, produced
a d i s t r i b u t e d e r f e c t on t h e Macn numDer a i d T Y D u i i u ~ ~ .
The power required f o r supersonic operation was i n s u b s t a n t i a l
sgreement with t h e smallest of several estimates obtained from
previously published information.
NACA RM L5OA03a
INTRODUCTION
Because of t h e s p e c i a l problems encountered i n f l i g h t near sonic
speeds, an urgent need e x i s t s f o r a d d i t i o n a l aerodynamic t e s t i n g
f a c i l i t i e s f o r use a t Mach numbers near u n i t y . As a means of providing
t h e needed f a c i l i t i e s , l a r g e subsonic wind t u n n e l s a l r e a d y o p e r a t i n g a t
Mach numbers s l i g h t l y l e s s t h a n u n i t y may be provided with supersonic
nozzles t o permit t e s t i n g a t Mach numbers s l i g h t l y g r e a t e r t h a n u n i t y .
The conversion of t h e Langley 8-foot high-speed t u n n e l t o t r a n s o n i c
o p e r a t i o n has been considered and one phase of t h e problem has been t h e
design and o p e r a t i o n of a f i x e d nozzle of a x i a l symmetry t o permit
t e s t i n g at high-subsonic Mach numbers and a t a supersonic Mach number
of 1 . 2 . The t u n n e l i s of t h e s i n g l e - r e t u r n , c l o s e d - t h r o a t type, of
c i r c u l a r c r o s s s e c t i o n throughout, and t h e low-speed p a r t of t h e r e t u r n
passage i s open t o atmospheric p r e s s u r e . I n s t a l l a t i o n of t h e nozzle a s
a p l a s t e r l i n e r i n t h e t h r o a t region of t h e t u n n e l was undertaken l a t e
i n 1947 and t h e g r e a t e r p o r t i o n of t h e flow surveys r e p o r t e d h e r e i n
were obtained during t h e e a r l y p a r t of 1948. This paper, which should
be of i n t e r e s t t o t h o s e concerned with t r a n s o n i c wind t u n n e l s , covers
t h e design and operating c h a r a c t e r i s t i c s of t h e nozzle.
NOZZU DESIGN
Preliminary Considerations
The design of t h e nozzle w a s d e c i s i v e l y influenced by t h e o r i g i n a l
t u n n e l geometry and by t h e power a v a i l a b l e . Experimental d a t a from
preliminary surveys with v a r i o u s flow-expansion l i n e r s i n t h e t h r o a t
of t h e t u n n e l i n d i c a t e d t h a t t h e power r e q u i r e d at Mach numbers a s
g r e a t a s 1 . 3 should not exceed t h e 16,000 horsepower a v a i l a b l e . I n
o r d e r t o a s s u r e adequate o p e r a t i n g power under a l l t e s t conditions, t h e
nozzle was t h e r e f o r e conservatively chosen t o produce a Mach number
of 1 . 2 .
A c i r c u l a r nozzle was s e l e c t e d r a t h e r t h a n t h e simpler twodimensional r e c t a n g u l a r nozzle mainly because i t s i n s t a l l a t i o n involved
l e s s modification t o t h e e x i s t i n g tunnel, but s e v e r a l o t h e r advantages
r e s u l t from t h e c i r c u l a r c r o s s s e c t i o n . The boundary-layer displacement
t h i c k n e s s , knowledge of which i s r e q u i r e d i n t h e nozzle design, i s more
r e l i a b l y estimated f o r t h e c i r c u l a r nozzle, because t h e boundary l a y e r
i s i n general uniformly d i s t r i b u t e d over t h e s u r f a c e , whereas f o r t h e
r e c t a n g u l a r nozzle t h e boundary l a y e r t e n d s t o t h i c k e n more r a p i d l y i n
t h e corners. Moreover, because of t h i s boundary-layer behavior and
a l s o because of t h e l e s s e r perimeter, with given c r o s s s e c t i o n , t h e
c i r c u l a r nozzle i s expected t o r e q u i r e l e s s power t h a n a r e c t a n g u l a r
NACA R h L5OA03a
3
nozzle with the same mass flow. The c i r c u l a r section may a l s o be
advantageous i n f a c i l i t a t i n g t h e t e s t i n g of p r o p e l l e r s and bodies of
revolution.
Several known disadvantages a r e associated with t h e c i r c u l a r cross
section. Perhaps t h e most serious of these i s t h e tendency of d i s turbances a r i s i n g from inaccuracies i n design o r from axisymmetrical
i r r e g u l a r i t i e s a t t h e wall t o become concentrated near t h e center of
t h e nozzle. Because of t h i s tendency, s p e c i a l care i s required t o
assure accuracy of design and i n s t a l l a t i o n . It i s believed e s s e n t i a l ,
f o r instance, t o allow f o r v a r i a t i o n s i n displacement thickness of t h e
boundary l a y e r . The c i r c u l a r cross section i s a l s o disadvantageous with
respect t o changing from one supersonic Mach number t o another (by
means of removable o r adjustable nozzle blocks) and with respect t o
providing windows f o r the use of flow-visualization equipment.
The shape of t h e p l a s t e r l i n e r was l a r g e l y d i c t a t e d by t h e o r i g i n a l
tunnel l i n e s . (see f i g . 1 . ) A long, slowly converging entrance l i n e r
was required i n order t o produce a steady uniform flow a t t h e t h r o a t
( ~ a c hnumber 1 . 0 ) , since such a uniform flow was t o be assumed i n t h e
t h e o r e t i c a l nozzle design. Inasmuch a s n e i t h e r t h e required precision
of t h i s flow nor t h e precision a t t a i n a b l e with a given geometrical
arrangement was known, t h e entrance l i n e r was made a s long and a s
slowly convergent a s the o r i g i n a l tunnel shape i n combination with
other requirements, such a s necessary length of supersonic portion and
required thickness of p l a s t e r , would allow. The length of t h e divergent
portion of t h e nozzle was chosen a s 90 inches, which i s about 1.5 times
the minimum possible length f o r a nozzle producing a Mach number of 1 . 2 .
This choice of length was believed t o be conservative with respect t o
t h e production of a s a t i s f a c t o r i l y uniform flow i n t h e t e s t region.
The thickness of t h e l i n e r , which a f f e c t s i t s f a i r i n g i n t o t h e o r i g i n a l
tunnel walls a t i t s upstream and downstream ends, was d i c t a t e d by t h e
requirement of a t l e a s t 314 inch of p l a s t e r t o permit proper anchoring
t o t h e o r i g i n a l s t e e l wall. I n meeting these various requirements, the
l i n e r extended 140 inches upstream and f a i r e d i n t o t h e o r i g i n a l tunnel
125 inches downstream from the t h r o a t . Space l i m i t a t i o n s r e s t r i c t e d
t h e t e s t section length t o only about one-half the tunnel diameter
unless s u f f i c i e n t power were available t o draw the shock i n t o the
d i f f u s e r . Lack of space would i n any case have prevented t h e i n s t a l l a t i o n of a second t h r o a t , but with Mach numbers not much g r e a t e r than 1.2
t h e e f f i c i e n c y of d i f f u s i o n through the normal shock i s believed t o be
l i t t l e i f any l e s s than t h a t i n a d i f f u s e r operating between t h e same
two Mach numbers. The requirement of a long, slowly converging
entrance l i n e r lends i t s e l f t o t h e production of a short subsonic t e s t
section near t h e t h r o a t , and such a t e s t section was provided.
(see
f i g . 1.) A narrow window i n t h e nozzle wall extending from upstream
of t h e t h r o a t t o downstream of t h e supersonic t e s t section was provided
f o r v i s u a l observation of flow phenomena.
NACA RM L50A03a
Potential-Flow Design
Divergent portion of nozzle.- The steady supersonic p o t e n t i a l flow
i n t h e a x i a l l y symmetric divergent portion of t h e nozzle was calculated
by t h e metho& of c h a r a c t e r i s t i c s . (see reference 1.) For t h e applicat i o n of t h i s method, a Mach number d i s t r i b u t i o n along the a x i a l c e n t e r
l i n e was a r b i t r a r i l y selected. (see f i g . 2 . ) The Mach number increased
from 1 . 0 a t t h e t h r o a t t o 1 . 2 a t point A 60 inches downstream of t h e
t h r o a t . The calculation then proceeded from t h e assumed values of Mach
number along t h e center l i n e and from t h e constant value 1 . 2 along t h e
c h a r a c t e r i s t i c through A, marking t h e upstream boundary of t h e uniformflow region.
The a x i a l Mach number d i s t r i b u t i o n was chosen with zero a x i a l Mach
number gradient at t h e t h r o a t i n order t o be consistent with t h e assumpt i o n of uniform flow a t t h a t p o i n t . The accuracy of t h e step-by-step
flow c a l c u l a t i o n s decreased near t h e t h r o a t and t h e Mach number
i n t e r v a l s between points of t h e Mach net were reduced. The f a c t t h a t
t h e chosen Mach number gradients were small i n the region near t h e
t h r o a t was considered favorable t o t h e accuracy of t h e calculation.
The c h a r a c t e r i s t i c network f o r t h e flow i n the divergent p o r t i o n
of t h e nozzle i s shown i n f i g u r e 3. The numerical calculations were
s t a r t e d a t point A and were extended s t e p by s t e p i n t o t h e remainder
of the f i e l d . Note t h a t the nozzle boundary was not immediately defined
but t h a t t h e network was extended beyond t h e probable position of t h e
w a l l s . This method of calculating from the center outward and from
l a r g e r t o smaller values of t h e Mach number i s believed t o be more
accurate t h a n t h e opposite method of calculation. I n p a r t i c u l a r ,
attempts t o calculate from a given wall shape t o t h e flow near t h e
center l e d t o d i f f i c u l t i e s due t o magnification of t h e inaccuracies
inherent i n a step-by-step computing process.
Once t h e flow f i e l d had been calculated, t h e Mach numbers and t h e
flow angles throughout the f i e l d were known. From t h e flow angles t h e
streamlines cquld be obtained by a method suggested by M r . Morton Cooper
of the Compressibility Division of t h e Langley Aeronautical Laboratory.
The streamline forming t h e e f f e c t i v e boundary of t h e nozzle was assumed
t o pass through t h e t h r o a t a t a radius equal t o the geometrical r a d i u s
of t h e tunnel decreased by t h e estimated displacement thickness
(about 0.23 i n . ) of t h e boundary l a y e r . The increment i n radius of t h i s
streamline a t downstream s t a t i o n s was obtained by i n t e g r a t i n g t h e
r= fl
e it d i d not
exceed 0.0175 radian) along a l i n e CD ( f i g . 3) p a r a l l e l t o t h e a x i s OA,
Special care i s required t o assure t h e accuracy of t h e i n t e g r a t i o n .
The angles used i n t h i s i n t e g r a t i o n should s t r i c t l y have been those
l y i n g along t h e streamline i t s e l f . A second approximation t o t h e t r u e
streamline was therefore obtained by taking t h e l i n e i n t e g r a l with
NACA RM L5OA03a
r e s p e c t t o d i s t a n c e p a r a l l e l t o OA of t h e flow a n g l e s along t h e approximate streamline CE obtained by t h e f i r s t i n t e g r a t i o n . This process i s
r a p i d l y convergent and a t h i r d approximation CB yielded no f u r t h e r
change i n r a d i i . The flow angles used i n t h e s e i n t e g r a t i o n s were c a r e f u l l y determined by i n t e r p o l a t i o n from t h e p o i n t s of t h e Mach n e t . The
e f f e c t i v e r a d i i ( t h a t i s , without adjustment f o r boundary-layer d i s placement) determined f o r t h e divergent p o r t i o n of t h e nozzle and f o r
t h e supersonic t e s t s e c t i o n a r e given i n t a b l e I .
One-dimensional theory was used t o check t h e o v e r - a l l expansion of
t h e e f f e c t i v e c r o s s - s e c t i o n a l a r e a of t h e nozzle from t h e minimum
s e c t i o n OC ( f i g . 3) where t h e flow i s one-dimensional t o t h e supersonic
t e s t s e c t i o n where t h e flow i s again one-dimensional, The o v e r - a l l
i n c r e a s e i n nozzle r a d i u s y i e l d e d by t h e one -dimensional t h e o r y d i f f e r e d
from t h a t obtained by t h e c h a r a c t e r i s t i c method by l e s s t h a n 0.001 inch.
Convergent p o r t i o n of nozzle.- The convergent p o r t i o n of t h e nozzle
was designed t o meet t h e requirement of i n c r e a s i n g l y gradual convergence
toward t h e t h r o a t , where t h e Mach number u n i t y is-rkached, and t o
produce a s h o r t subsonic t e s t s e c t i o n i n t h e v i c i n i t y of t h e t h r o a t .
With i n c r e a s e i n Mach number toward u n i t y , t h e flow becomes i n c r e a s i n g l y
s e n s i t i v e both t o changes i n c r o s s s e c t i o n of t h e stream tube and t o t h e
curvature of t h e s u r f a c e . The entrance l i n e r was t h e r e f o r e desigped
with t h e curvature i n t h e d i r e c t i o n of flow decreasing t o zero a t t h e
throat.
The s e n s i t i v i t y of t h e flow i s such t h a t considerable care was
r e q u i r e d t o a s s u r e t h a t sonic speed would a c t u a l l y be a t t a i n e d a t t h e
assumed t h r o a t . Adjustment t o allow f o r development of t h e boundary
l a y e r was t h e r e f o r e r e q u i r e d . With unif o m flow ( zero pressure g r a d i e n t )
t h e slope of t h e boundary l a y e r was estimated a t 0.0018. I f , however,
t h e v e l o c i t y i s i n c r e a s i n g through t h e t h r o a t , t h e boundary-layer
t h i c k n e s s w i l l i n c r e a s e l e s s r a p i d l y . I n fac-t, it seems q u i t e p o s s i b l e
t h a t a t Mach number u n i t y on t h e supersonic s i d e a zero-gradient flow
would be u n s t a b l e . Any a c c i d e n t a l t h i n n i n g of t h e boundary l a y e r would
t h e n produce an increment i n Mach number, and t h e accompanying p r e s s u r e
g r a d i e n t would produce t h e boundary-layer t h i n n i n g required t o m a i n t a i n
t h e Mach number g r a d i e n t . Moreover, t h i s e f f e c t might be progressive
so t h a t t h e e f f e c t i v e t h r o a t would move upstream. I n order t o i n s u r e
s t a b i l i t y of t h e e f f e c t i v e t h r o a t p o s i t i o n t h e slope of t h e entrance
l i n e r a t t h e t h r o a t was taken a s 0.0015 i n s t e a d of t h e value 0.0018
estimated from t h e boundary-layer growth.
>
curvature, and slowness of convergence and a l s o connects reasonably
smoothly with t h e o r i g i n a l t u n n e l l i n e s gives t h e geometric ~ a d i u si n
inches a s
NACA RM L50A03a
where 46.25 i s the geometric radius a t t h e e f f e c t i v e minimum section
i n inches ( f i g . 1) and x i s t h e distance i n inches upstream of t h e
e f f e c t i v e minimum section. This expression was used f o r calculating
t h e r a d i i given i n t a b l e 11. The geometric minimum occurs 30 inches
upstream of t h e e f f e c t i v e minimum s e c t ion.
Boundary-Layer Compensation
The necessity f o r taking account of t h e boundary-layer development
has already been pointed out i n connection with t h e design of t h e
entrance l i n e r . With supersonic Mach numbers a s low a s t h a t f o r which
t h i s nozzle i s designed it i s a l s o imperative t o allow f o r t h e boundaryl a y e r development i n t h e divergent p a r t of t h e nozzle.
In t h e absence of experimental data on t h e behavior of t h e turbul e n t boundary l a y e r i n transonic flow, c e r t a i n reasonable assumptions
were made. The boundary-layer behavior was assumed t o be e s s e n t i a l l y
t h e sane, except f o r d i r e c t e f f e c t s of density changes, a s f o r incomp r e s s i b l e flow; and, j u s t a s with incompressible flow, t h e flow outside
t h e boundary l a y e r was assumed t o behave a s i f t h e streamlines near t h e
boundary were displaced away from t h e wall by an amount equal t o t h e
displacement thickness of the boundary l a y e r .
The calculation of displacement-thickness d i s t r i b u t i o n s i s made
through use of t h e boundary-layer momentum equation. Through t h i s
equation the influence of the outside flow upon the development of t h e
boundary l a y e r i s determined. Approximate formulas f o r t h e computation
of turbulent-boundary-layer momentum thicknesses i n compressible flows
a r e given i n reference 2 . The calculations require a knowledge of t h e
Mach and Reynolds number d i s t r i b u t i o n s along t h e tunnel, t h e i n i t i a l
boundary-layer condition, and t h e v a r i a t i o n of an important boundaryl a y e r parameter Hc, t h e r a t i o of t h e displacement thickness t o t h e
momentum thickness. The displacement thickness i s obtained from t h e
momentum thickness by m u l t i p l i c a t i o n with Hc. The v a r i a t i o n of Hc
was determined experimentally-from data obtained i n the Langley 8-foot
high-speed tunnel during a preliminary nozzle investigation i n which
rough wooden nozzles were t e s t e d i n the tunnel. For the conditions
e x i s t i n g i n t h e 8-foot tunnel, Rc was found t o vary mainly with Mach
number and t o follow approximately the v a r i a t i o n suggested i n r e f e r ence 2 . A more extended discussion of t h i s v a r i a t i o n i s given i n r e f e r ence 3.
The nozzle design was f a c i l i t a t e d by taking boundary-layer measurements along t h e wall of t h e o r i g i n a l tunnel. A preliminary check of t h e
NACA RM L5OA03a
a p p l i c a b i l i t y of t h e calculation methods of reference 2 t o the predict i o n of the boundary-layer development a x i a l l y along the wall of t h e
8-foot tunnel was made by calculating the displacement-thickness growth
i n a subsonic l i n e r operating a t a t e s t - s e c t i o n Mach number of 0.85 and
comparing the calculated values with measured values of t h e bomdaryl a y e r displacement thickness a t t h e tunnel w a l l ( f i g . 4 ) . I n t h i s
preliminary calculation t h e v a r i a t i o n of Hc with Mach number was
neglected and a constant value of 1.28 was used. This procedure
involved only small e r r o r since t h e v a r i a t i o n of Hc with Mach number
up t o a Mach number of 0.85 was not very g r e a t . The agreement between
theory and experiment ( f i g . 4) was considered s a t i s f a c t o r y .
The c a l c u l a t i o n of t h e d i s t r i b u t i o n of boundary-layer displacement
thickness f o r use i n the nozzle design proceeded from a measured value
f a r upstream i n t h e contraction cone. (see f i g . 4 . ) I n t h e upstream
portion of t h e nozzle, t h e Mach number d i s t r i b u t i o n needed i n t h e calcul a t i o n was determined from the cross-sectional area a t each point i n
connection with t h e assumption of uniform a x i a l (one-dimensional) flow
a t each section and Mach number u n i t y a t t h e t h r o a t . For such a slowly
converging entrance, the assumption of one-dimensional flow involves
negligible e r r o r . I n estimating the Mach numbers i n the s e n s i t i v e
region near the t h r o a t , t h e previously estimated slope of t h e displacement thickness was taken i n t o account. I n t h e divergent portion of the
nozzle, t h e Mach number d i s t r i b u t i o n a t the wall was taken from t h e
potential-flow design. The calculated Mach number d i s t r i b u t i o n i s shown
i n f i g u r e 5 . The v a r i a t i o n with Mach number of the r a t i o Ec was taken
from the r e s u l t s of the preliminary investigations.
The calculated course of t h e displacement thickness i s shown i n
f i g u r e 6. Note t h a t a decrease i n thickness i s predicted i n t h e region
of r a p i d a c c e l e r a t i o n just downstream from the t h r o a t . The experimental
data presented i n f i g u r e 6 a r e discussed i n a subsequent section i n
connection with t h e r e s u l t s of flow surveys i n t h e nozzle. The geometric
radius a t any section of t h e divergent p a r t of the nozzle i s obtained by
adding t h e boundary-layer displacement thickness a t t h a t section t o t h e
radius determined i n the portential-flow design. The geometric r a d i i so
obtained a r e presented i n t a b l e I.
NOZZM INSTALLATION
I n s t a l l a t i o n of the nozzle a s a p l a s t e r l i n e r inside t h e o r i g i n a l
walls of t h e tunnel involved t h e use of s p e c i a l construction materials
and methods. A high-strength, quick-setting p l a s t e r , Hydrocal B-11, was
selected f o r the construction material a f t e r p r a c t i c a l t e s t s indicated
t h a t it was s a t i s f a c t o r y f o r t h e purpose of t h e temporary i n s t a l l a t i o n .
The p l a s t e r was very hard and r e s i s t a n t t o chipping and on s e t t i n g
increased only s l i g h t l y i n volume.
The method used f o r i n s t a l l i n g t h e p l a s t e r l i n e r c o n s i s t e d
e s s e n t i a l l y of f a s t e n i n g metal l a t h t o t h e s t e e l w a l l s of t h e t u n n e l
(by welding) i n t h e region s e l e c t e d f o r t h e l i n e r i n s t a l l a t i o n and
applying p l a s t e r coats t o b u i l d up t h e l i n e r t o t h e proper geometric
dimensions. ( s e e t a b l e s I and 11.) A l l l e a k s i n t h e t u n n e l w a l l i n
t h e r e g i o n designated f o r t h e l i n e r i n s t a l l a t i o n were stopped (by
welding) before t h e p l a s t e r was a p p l i e d . The f i n a l p l a s t e r coat was
shaped t o t h e d e s i r e d a x i a l p r o f i l e by means of a template r i g which
was designed t o r i d e on metal r i n g s f a s t e n e d s e c u r e l y t o t h e t u n n e l
w a l l a t s t a t i o n s 2 f e e t a p a r t a x i a l l y along t h e nozzle l e n g t h . The
reference r i n g s were ground a s n e a r l y a s p o s s i b l e t o t r u e c i r c u l a r
shape by means of a grinding attachment r o t a t i n g off a r i g i d tube
a l i n e d along t h e t u n n e l a x i s of symmetry. The average r a d i i of t h e
v a r i o u s reference r i n g s were determined very a c c u r a t e l y , u s i n g t h e
c e n t r a l tube a s a reference, and no d e v i a t i o n s of more t h a n 0.004 inch
from t r u e c i r c u l a r shape were obtained. This technique of b u i l d i n g up
t h e p l a s t e r l i n e r l e f t a t t h e r i n g l o c a t i o n s narrow u n f i l l e d channels
which r e q u i r e d f i l l i n g and sanding a f t e r t h e r e s t of t h e nozzle was
i n s t a l l e d . F i l l i n g was a l s o r e q u i r e d a t t h e edges of a long narrow
window i n s t a l l e d i n t h e t o p of t h e t u n n e l f o r observation of flow
phenomena.
A d e t a i l view showing s t a g e s of t h e p l a s t e r - a p p l i c a t i o n technique
i s given i n f i g u r e 7. I n t h e foreground of t h i s view, metal l a t h i s
shown a t t a c h e d t o t h e t u n n e l wall; i n t h e c e n t e r of t h e view, a b a s i c
coat of p l a s t e r i s shown a p p l i e d t o t h e metal l a t h ; and i n t h e background, a s t r i p of t h e f i n a l smooth p l a s t e r coat i s v i s i b l e . I n
f i g u r e 8, t h e template r i g used f o r shaping t h e f i n a l p l a s t e r coat t o
t h e d e s i r e d p r o f i l e between adjoining reference r i n g s i s shown r e s t i n g
on t h e t u n n e l f l o o r . A photograph of t h e completed nozzle, a s viewed
from downstream, i s given as f i g u r e 9.
The accuracy of t h e i n s t a l l a t i o n was of importance p r i m a r i l y i n
t h e t h r o a t and divergent p o r t i o n of t h e nozzle. Physical measurement
of t h e o v e r - a l l accuracy of t h e nozzle i n s t a l l a t i o n was not f e a s i b l e
because of t h e l a r g e amount of time and c a r e f u l work r e q u i r e d f o r such
an undertaking. The a x i a l p r o f i l e of t h e nozzle w a l l w a s c a r e f u l l y
checked, however, by means of templates of t h e design geometric shape
extending from 24 inches upstream of t h e e f f e c t i v e minimum t o 96 inches
downstream of t h e e f f e c t i v e minimum s e c t i o n . These checks of t h e a x i a l
p r o f i l e by means of templates were made a t 2 -inch i n t e r v a l s around t h e
e n t i r e circumference of t h e nozzle, and maxi
w a l l shape from t h e design shape d i d not exc
maximum deviations i n t h e o r i g i n a l i n s t a l l a t i o n were subsequently
reduced (by sanding and f i l l i n g ) t o d e v i a t i o n s of t h e order of 0.003 inch.
The f i n i s h e d wall surface, although glazed i n appearance and smooth t o
t h e touch (except f o r s l i g h t d i s c o n t i n u i t i e s i n slope a t some r e f e r e n c e r i n g l o c a t i o n s ) was not so f i n e a s t h a t comonly used f o r small
NACA RM L5OA03a
supersonic nozzles.
was very smooth.
Relative t o t h e tunnel s i z e , however, t h e surface
SURVEYS OF NOZZLF: FLOW
Apparatus and Measurements
The arrangement shown i n f i g u r e 10 was used f o r t h e measurement of
s t a t i c pressures a t t h e wall and near t h e a x i a l center l i n e of t h e
nozzle. The s t a t i c - p r e s s u r e o r i f i c e s i n the surfaces of both t h e w a l l
and t h e 2-inch-diameter c y l i n d r i c a l axial-survey tube were of 0.031-inch
diameter and were i n s t a l l e d normal t o t h e surface. Wall o r i f i c e s were
located a x i a l l y 2 inches a p a r t i n t h e t h r o a t and supersonic-flow region
and 6 inches a p a r t i n other regions of t h e nozzle. Those i n the
c y l i n d r i c a l tube were located 2 inches a p a r t i n t h e t h r o a t and supersonic t e s t section and 6 inches a p a r t elsewhere. Wall o r i f i c e s were
i n s t a l l e d 90' a p a r t around t h e circumference of t h e channel a t several
a x i a l s t a t i o n s t o permit checks f o r symmetry of the flow. Wall and
cylindrical-tube surfaces near s t a t i c - p r e s s u r e o r i f i c e s were kept f r e e
of i r r e g u l a r i t i e s
Local s t a t i c -pressure measurements by means of
o r i f i c e s i n t h e cylindrical-tube and nozzle-wall surfaces were assumed
t o be equal t o those outside t h e boundary l a y e r except i n t h e region of
shock where pressure changes would occur over an a x i a l distance g r e a t e r
a t the surface than outside t h e boundary l a y e r .
.
Total-pressure measurements i n the subsonic flow upstream of the
nozzle t h r o a t were obtained by means of a total-pressure tube i n t h e
e l l i p s o i d a l nose a t t h e upstream end of t h e c y l i n d r i c a l tube. ( s e e
f i g . 1 0 . ) The t o t a l pressures were assumed t o be correct a s measured
and t o apply downstream of t h e s t a t i o n of measurement a s long a s t h e
flow remained e s s e n t i a l l y shock f r e e .
The c y l i n d r i c a l tube was positioned along t h e a x i a l center l i n e of
t h e nozzle so t h a t t h e upstream end of t h e tube was located s u f f i c i e n t l y
f a r upstream of t h e t h r o a t t o introduce no appreciable disturbances i n
the flow near the e f f e c t i v e minimum section. The tube was maintained
i n p o s i t i o n by a r i g i d support system located i n t h e tunnel d i f f u s e r
( f i g . 10) and by sweptback s t a y wires running from t h e wall of t h e cont r a c t i o n cone t o t h e nose of the tube. The tube was capable of a x i a l
adjustment t o permit s t a t i c - p r e s s u r e measurements a t i n t e r v a l s a s close
a s desired.
Cones (of 3' included angle ) equipped with s t a t i c -pre ssure o r i f i c e s
and t o t a l - p r e s s u r e tubes were used f o r more d e t a i l e d surveys of t h e flow
i n the supersonic t e s t section than could be obtained by use of t h e
c y l i n d r i c a l tube. Static-pressure o r i f i c e s were of 0.010-inch diameter
10
NACA RM L5OA03a
and were i n s t a l l e d normal t o t h e surface a t 2-inch i n t e r v a l s along the
length of t h e cone. O r i f i c e s were a l s o i n s t a l l e d a t angular locat i o n s 90' a p a r t i n t h e cone surfaces t o permit approximate checks f o r
symmetry and i n c l i n a t i o n of t h e flow. Total-pressure tubes of
0.010-inch inside diameter were i n s t a l l e d i n cone t i p s which were
interchangeable with sharp t i p s normally used with the cones. The cones
were capable of a x i a l adjustment t o permit measurements a t a x i a l
i n t e r v a l s a s close as desired. S t a t i c pressures measured by means of
t h e cones required correction f o r induced v e l o c i t y a t t h e surfaces of
t h e cones; but t h i s correction was very s m a l l , corresponding t o a
t h e o r e t i c a l l y estimated Mach number increment of the order of 0.001 and
was near t h e accuracy of t h e pressure measurements. A l a r g e cone
66 inches long ( f i g . 11) was designed f o r flow surveys very near t h e
a x i a l center l i n e of t h e supersonic t e s t section, and an arrangement of
two small cones located 180° a p a r t angularly ( f i g . 12) was used f o r
surveys a t distances of 3, 7, and 13 inches off t h e center l i n e .
S a t i s f a c t o r y precision i n pressure measurements was more d i f f i c u l t
t o a t t a i n by means of o r i f i c e s i n cones than by means of o r i f i c e s i n
t h e c y l i n d r i c a l tube, because t h e t h i n n e r boundary l a y e r on t h e cones
rendered t h e cone surface conditions more c r i t i c a l . The t h i n boundary
l a y e r s on t h e cones were considered advantageous, however, f o r pressure
measurements i n t h e v i c i n i t y of shocks.
Pressures were measured with multiple-tube manometers containing
tetrabromoethane. These manometers were photographed simultaneously.
The random e r r o r i n t h e pressure measurements was estimated t o be no
g r e a t e r than 3 pounds per square f o o t , which i s equivalent t o an e r r o r
of l e s s than 0.0033 i n the flow Mach number throughout a Mach number
range extending from 0.4 t o 1 . 3 .
A shadow system was provided t o supplement the pressure apparatus
i n examination of t h e supersonic flow f o r t h e presence of s$.rong d i s turbances. The equipment used f o r producing shadow images due t o
changes i n density gradients i n t h e supersonic flow consisted of an
intense point-source l i g h t and a condensing l e n s which were used t o
p r o j e c t a beam of approximately p a r a l l e l l i g h t rays across t h e nozzle
diameter. The shadow equipment w a s made portable f o r g r e a t e r v e r s a t i l i t y
i n examination of flow disturbances. Disturbance images were observed
on t h e nozzle w a l l opposite t h e observation window. The s e n s i t i v i t y of
t h e system was s u f f i c i e n t t o permit the observation of normal shocks a t
Mach numbers a s low a s 1.06.
The stagnation temperature of t h e flow mixture i n t h e tunnel was
measured by means of e l e c t r i c a l - r e s i s t a n c e thermometers located a t
several points between t h e t u ~ n e wall
l
and center l i n e i n t h e low-speed
region upstream of t h e entrance cone. The s t a t i c temperature, equivalent
NACA RM L5OA03a
11
t o t h e stagnation temperature in t h e low-speed region, w a s assumed t o
decrease I s e n t r o p i c a l l y with flow expansion.
Apparatus used f o r measurement of pressures and temperature i n t h e
boundary l a y e r adjacent t o t h e nozzle w a l l and methods used f o r reduct i o n of t h e measurements a r e described i n reference 3.
The measurements reported i n t h i s paper were, with t h e exception
of those made during several. runs i n which condensation e f f e c t s were
d e l i b e r a t e l y introduced, obtained with s u f f i c i e n t l y high tunnel temperat u r e s t o preclude t h e existence of s i g n i f i c a n t condensation e f f e c t s i n
t h e flow. For t y p i c a l tunnel-operation temperatures, t h e Reynolds
numbers a t t a i n e d i n t h e Mach number 1.2 t e s t section were approximately 3.8 x 106 per f o o t . Mach numbers were obtained from t h e r a t i o
of s t a t i c t o t o t d pressures.
Results and Discussion
Mach number d i s t r i b u t i o n s . - Mach number d i s t r i b u t i o n s obtained
from pressure measurements a t t h e wall and at t h e surface of t h e
c y l i n d r i c a l tube along t h e center l i n e of t h e nozzle f o r both subsonic
and supersonic operation a r e presented i n f i g u r e s 13 and 14. These
r e s u l t s indicated t h a t with subsonic operation t h e flow i n t h e t h r o a t
region of t h e nozzle was very uniform and s u i t a b l e f o r subsonic t e s t i n g
purposes. An e s s e n t i a l l y zero-gradient region e x i s t e d i n t h e flow over
a considerable length i n t h e v i c i n i t y of t h e t h r o a t f o r Mach numbers
up t o 0.97, but above 0.97 a s t h e Mach number approached u n i t y t h e
gradients i n t h e downstream p o r t i o n of t h i s region gradually increased.
(see f i g . 13. ) The flow i n t h e upstream portion of t h i s subsonic t e s t
s e c t i o n was uniform up t o Mach numbers a s high a s 0.99; and a t t h i s
1
Mach number a model 8 inches long and 1- inches i n diameter was success3
f u l l y t e s t e d with negligible interference from t h e tunnel walls. The
length of t h e uniform t e s t region was such a t any Mach number t h a t with
fineness r a t i o s about 6 t h e s i z e of t h e model %hat could be t e s t e d was
l i m i t e d by choking r a t h e r than by t h e length of the t e s t section.
With supersonic operation t h e experimental Mach number d i s t r i b u t i o n
i s seen t o be i n excellent agreement with t h e design d i s t r i b u t i o n
( f i g s . 13 and 1 4 ) . The f a c t t h a t t h e experimental p o s i t i o n of t h e
t h r o a t ( ~ a c hnumber u n i t y ) i s within an inch of t h e design p o s i t i o n i s
regarded a s p a r t i c u l a r l y s i g n i f i c a n t . A comparison of t h e a c t u a l
d i s t r i b u t i o n obtained i n t h e contraction cone or convergent portion
of t h e nozzle with one-dimensional theory, which includes calculated
boundary-layer-displacement e f f e c t s , i s presented i n f i g u r e 15. Near
t h e t h r o a t region of the nozzle t h e agreement between t h e t h e o r e t i c a l
12
NACA RM L5OA03a
Mach number d i s t r i b u t i o n and t h a t measured along t h e contraction-cone
c e n t e r l i n e and w a l l was e s p e c i a l l y good. This agreement i n d i c a t e s
t h a t t h e flow a t t a i n e d near t h e minimum s e c t i o n was p r a c t i c a l l y onedimensional and f r e e of appreciable induced v e l o c i t i e s . The agreement
became i n c r e a s i n g l y poor w i t h d i s t a n c e upstream of t h e t h r o a t r e g i o n
l a r g e l y because of induced v e l o c i t i e s b r o u a t about by i n c r e a s i n g l y
r a p i d changes of t h e contraction-cone r a d i i . The g e n e r a l l y f a v o r a b l e
agreement between t h e o r y and experiment i n t h e c o n t r a c t i o n cone appeared
t o v a l i d a t e t h e use of one-dimensional flow r e l a t i o n s i n t h e d e s i g n of
g r a d u a l l y converging channels f o r subsonic flow.
I n order t o o b t a i n a more a c c u r a t e i n d i c a t i o n of t h e c h a r a c t e r of
1
t h e flow, p r e s s u r e measurements were taken a t z-inch a x i a l increments
near t h e c e n t e r l i n e of t h e divergent p a r t of t h e nozzle. The c o r r e sponding Mach number d i s t r i b u t i o n i s presented i n f i g u r e 16. These
measurements were obtained by moving t h e c y l i n d r i c a l survey tube
( f i g . 10) a x i a l l y at
4- i n c h
increments between runs f o r e i g h t successive
runs. The flow near t h e a x i s i n t h e divergent p o r t i o n of t h e nozzle and
i n t h e supersonic t e s t s e c t i o n i s seen t o be f r e e of l a r g e di'sturbances.
The maximum flow d i s t u r b a n c e s i n t h e supersonic t e s t s e c t i o n were
equivalent t o d e v i a t i o n s of 0.02 from t h e design Mach number of 1 . 2 .
The Mach number v a r i a t i o n with d i s t a n c e o f f t h e c e n t e r l i n e , and
p a r t i c u l a r l y t h e tendency of t h e flow disturbances t o become concent r a t e d at t h e c e n t e r , w a s i n v e s t i g a t e d by means of t h e survey cones
( f i g s . 11 and 1 2 ) . The l a r g e r cone ( f i g . 11) could be moved p a r a l l e l t o
t h e nozzle a x i s t o p l a c e p r e s s u r e o r i f i c e s from 0 . 1 inch t o 1 . 5 inches
off t h e a x i s a t any s t a t i o n . Mach numbers obtained with t h i s cone i n
two p o s i t i o n s a r e compared i n f i g u r e 17 with those obtained from t h e
c y l i n d r i c a l tube ( o r i f i c e s 1 i n . o f f c e n t e r l i n e ) . The comparison i s
a f f e c t e d by t h e l a c k of p r e c i s i o n of t h e cone measurements and by a
change, with time between t e s t s , i n t h e disturbance a t t h e 77-inch s t a t i o n ; but w i t h i n t h e accuracy of t h e d a t a no c o n s i s t e n t v a r i a t i o n i n
t h e Mach number within 1.5 inches of t h e c e n t e r l i n e can be d e t e c t e d .
With t h e cone mounting arrangement of f i g u r e 12, t h e Mach number
d i s t r i b u t i o n could be obtained a t 3, 7, and 1 3 inches o f f t h e c e n t e r
l i n e . This arrangement w a s used t o t r a c e t h e extension from t h e c e n t e r
l i n e outward of disturbances near t h e 75-inch s t a t i o n . I n t h e two cases
i n v e s t i g a t e d (two i n t e n s i t i e s of t h e disturbance a s shown i n f i g s . 18
and 1 9 ) t h e disturbance i s seen t o decrease, a s expected, from t h e
c e n t e r outward. Considerable s c a t t e r p r e s e n t i n t h e f i r s t of t h e s e
cone measurem~ntswas reduced (downstream of t h e 7 3 . 5 - i n . s t a t i o n a t
7 i n . off t h e c e n t e r l i n e , f i g . 1 9 ) by improvement of t h e cone s u r f a c e .
These d i s t u r b a n c e s followed t h e Mach l i n e s of t h e flow and became
broader outward from t h e c e n t e r l i n e , and because of t h i s reason a s
NACA RM L5OA03a
w e l l as because of t h e tendency t o become concentrated a t t h e c e n t e r ,
t h e y a r e believed t o be due not t o shock waves but t o gradual convergence of Mach l i n e s of compression. The Mach number d i s t r i b u t i o n s
a t 7 and 1 3 inches off t h e c e n t e r l i n e may be compared i n f i g u r e 18
with those c a l c u l a t e d by t h e c h a r a c t e r i s t i c s method s t a r t i n g from t h e
experimental d i s t r i b u t i o n a t 1 inch o f f t h e c e n t e r l i n e . From t h e s e
i n v e s t i g a t i o n s , t h e flow appears t o be more uniform elsewhere t h a n a t
the center.
From t h e s e surveys and from shadowgraph observations, t h e flow i n
t h e t e s t s e c t i o n appears t o be f a i r l y uniform and f r e e of shocks, a
conclusion which i s f u r t h e r supported by t h e agreement between t h e
experimental and design Mach number d i s t r i b u t i o n s . I f g r e a t e r
uniformity of t h e Mach number d i s t r i b u t i o n t h a n t h a t e x i s t i n g near t h e
c e n t e r l i n e i s r e q u i r e d f o r any p a r t i c u l a r t e s t , t h e model can be
placed i n an o f f - c e n t e r p o s i t i o n . By t h e procedures hereinbefore
described, a c i r c u l a r nozzle of s i z e comparable t o t h e 8-foot high-speed
t u n n e l supersonic nozzle f o r Mach number 1 . 2 can e v i d e n t l y be designed
t o produce a s a t i s f a c t o r i l y uniform supersonic flow.
E f f e c t s of w a l l i r r e g u l a r i t i e s on flow uniformity.- The supersonic
f l o w i n t h e nozzle immediately a f t e r i n s t a l l a t i o n was not so smooth a s
d e s i r e d ( c i r c u l a r symbols, f i g . 20) and a l i m i t e d amount of time was
spent i n attempting t o improve t h e flow. It w a s found t h a t t h e flow
d e v i a t i o n s measured near t h e nozzle c e n t e r l i n e could be t r a c e d t o
i r r e g u l a r i t i e s i n t h e wall shape by use of t h e c a l c u l a t e d c h a r a c t e r i s t i c s network ( f i g . 3 ) . The use of t h i s network f o r l o c a t i n g t h e
p o i n t s of o r i g i n of flow d i s t u r b a n c e s w a s made p o s s i b l e by t h e c l o s e
agreement of t h e a c t u a l and design flows. Careful checks of t h e w a l l
shape by means of a x i a l - p r o f i l e templates i n regions where flow d i s turbances were suspected t o o r i g i n a t e i n d i c a t e d small d e v i a t i o n s from
t h e design shape. These d e v i a t i o n s u s u a l l y c o n s i s t e d of shallow bumps
and depressions, sometimes axisymmetrical and sometimes l o c a l i z e d , i n
t h e nozzle wall. I n each of s e v e r a l i n s t a n c e s where w a l l i r r e g u l a r i t i e s
were reduced by sanding bumps and f i l l i n g depressions, t h e a s s o c i a t e d
flow d i s t u r b a n c e s were observed t o have decreased. Deviations of t h e
a c t u a l nozzle p r o f i l e from t h e design p r o f i l e , a s determined by means
of a x i a l - p r o f i l e templates, were reduced from rt0.011 inch t o 90.005 inch
i n a region extending from 24 inches upstream of t h e e f f e c t i v e minimum
t o 96 inches downstream of t h e e f f e c t i v e minimum s e c t i o n . Figure 2G
shows a comparison of t h e measured Mach number d i s t r i b u t i o n s a x i a l l y
along t h e w a l l and near t h e c e n t e r l i n e of t h e nozzle before and a f t e r
r e d u c t i o n of t h e wall-prof i l e d e v i a t i o n s ; a n o t i c e a b l e improvement i n
t h e flow i s evident, e s p e c i a l l y near t h e c e n t e r l i n e a t a s t a t i o n
75 inches downstream of t h e e f f e c t i v e minimum s e c t i o n where t h e deviat i o n from t h e design Mach number ( 1 . 2 ) was reduced t o about one-half
of t h e o r i g i n a l d e v i a t i o n . Further improvement of t h e flow was cons i d e r e d p o s s i b l e by a d d i t i o n a l work on t h e i n s t a l l a t i o n accuracy,
14
NACA RM L5OA03a
although exact agreement of a c t u a l and design flow d i s t r i b u t i o n s could
not be expected because of p o s s i b l e e r r o r i n t h e boundary-layer p r e d i c t i o n and because of p o s s i b l e p e r i o d i c flow disturbances not due t o
w a l l irregularities
.
An i n d i c a t i o n of t h e r e l a t i v e l y small bump dimensions r e q u i r e d t o
produce severe flow disturbances was obtained by f a s t e n i n g a 0.030-inchdiameter s t r i n g around t h e nozzle w a l l a t an a x i a l s t a t i o n 30 inches
downstream of t h e e f f e c t i v e minimum s e c t i o n and observing i t s e f f e c t on
t h e ~ u p e r s o n i cflow. A strong compression disturbance followed by a
s t r o n g expansion ( f i g . 2 1 ) , with d e v i a t i o n s of a s much a s 0.13 from t h e
disturbance-free flow of Mach number 1.2, was produced near t h e nozzle
a x i s . The disturbance, which was strong near t h e nozzle a x i s , w a s
observed a s a weak compression and accompanying expansion at i t s i n t e r s e c t i o n with t h e w a l l i n t h e downstream p a r t of t h e supersonic t e s t
secti-on. (See f i g . 21.) Small i r r e g u l a r i t i e s of t h e nature of roughness, cracks i n t h e p l a s t e r surface, and small l o c a l d i s c o n t i n u i t i e s i n
slope produced no n o t i c e a b l e e f f e c t on t h e flow. This behavior i s
believed t o be p a r t i a l l y due t o t h e masking e f f e c t of t h e t h i c k boundary
l a y e r and p a r t i a l l y due t o t h e randorn n a t u r e of t h e s e disturbances, a s
c o n t r a s t e d w i t h t h e a x i a l l y symmetrical disturbance produced by t h e
string.
,
Boundary-layer development.- An attempt was made by means of
boundary-layer surveys t o check t h e c a l c u l a t e d displacement t h i c k n e s s e s
a t various p o s i t i o n s along t h e nozzle w a l l ( f i g . 6 ) . Although i n some
cases t h e experimentally determined values were i n good agreement with
t h e c a l c u l a t e d values, i n o t h e r s considerable divergence occurred.
Moreover, values of t h e displacement t h i c k n e s s obtained a t t h e same
time and a t t h e same a x i a l p o s i t i o n b u t a t d i f f e r e n t angular p o s i t i o n s
cln t h e w a l l of t h e nozzle f a i l e d t o agree among themselves. S e v e r a l
p o s s i b l e reasons f o r t h e s e divergence-s include l o c a l accumulations of
boundary-layer a i r due t o l a c k of p e r f e c t symmetry i n t h e p r e s s u r e
d i s t r i b u t i o n s , l o c a l i z e d l e a k s i n t o t h e t u n n e l upstream of t h e measuring
p o s i t i o n , and v a r i a t i o n of s t a g n a t i o n temperature (assumed c o n s t a n t )
through t h e boundary l a y e r . An a n a l y s i s of t h e boundary-layer v e l o c i t y
p r o f i l e s obtained i n t h i s i n v e s t i g a t i o n i s contained i n reference 3.
The success of t h e o v e r - a l l design i n producing t h e design Mach
number d i s t r i b u t i o n s , p a r t i c u l a r l y i n t h e v i c i n i t y of t h e t h r o a t ,
i n d i c a t e s t h a t t h e assumptions made i n t h e design a s t o t h e behavior of
t h e t u r b u l e n t boundary l a y e r i n t r a n s o n i c flow and a s t o i t s e f f e c t on
t h e o u t s i d e flow a r e a t l e a s t approximately c o r r e c t . I f t h e changes i n
boundary-layer displacement thickness had been neglected i n t h e design,
t h e nozzle would have been expected t o produce a Mach number of about 1.18
i n s t e a d of 1 . 2 0 .
NACA RM L50AO3a
OPERATIONAL CHARACTERISTICS
E f f e c t s of Humidity and Heat Transfer
Because t h e low-speed p a r t of t h e t u n n e l was open t o t h e atmosphere, care was r e q u i r e d t o prevent flow disturbances due t o condensat i o n i n t h e high-speed, low-temperature p a r t of t h e nozzle. Adverse
condensation e f f e c t s were prevented by allowing t h e t u n n e l t o become
heated, a method t h a t was q u i t e e f f e c t i v e i n winter when s t a g n a t i o n
temperatures not above 180' F were r e q u i r e d but was somewhat inconvenient i n summer when t h e r e q u i r e d s t a g n a t i o n temperatures might
reach 230' F. The s t a g n a t i o n temperature was c o n t r o l l e d by a d j u s t i n g
annulus around t h e t u n n e l i n t h e
t h e amount of a i r exchange through
low-speed s e c t i o n ,
Condensat i o n shocks, which have been t h e source of some d i f f i c u l t y
i n o t h e r supersonic tunnels, could be prevented according t o r e f e r ence 4 by l i m i t i n g supercooling t o l e s s t h a n 54' F. I n t h e supersonic
nozzle h e r e i n described, however, no condensation shock could be
d e t e c t e d e i t h e r from p r e s s u r e d i s t r i b u t i o n s o r from shadowgraph observat i o n s , even though a t times t h e nominal supercoolihg was allowed t o
exceed 54' F. On t h e o t h e r hand, f o g could u s u a l l y be observed i n t h e
nozzle. The f o g was most evident i n t h e cooler region near t h e w a l l
where, because of t h e incomplete mixing of t h e cooling air with t h e
remainder of t h e a i r i n t h e tunnel, t h e temperature might be a s much
a s 50' F l e s s t h a n t h a t a t t h e c e n t e r . The reason f o r t h i s behavior i s
believed t o be t h a t i n a continuous-circuit t u n n e l such a s t h e Langley
8-foot high-speed t u n n e l t h e s o l i d o r l i q u i d p a r t i c l e s , and p o s s i b l y
ions, c a r r i e d around i n t h e a i r stream a r e s u f f i c i e n t l y numerous t o
provide n u c l e i f o r condensation. Moreover, i n such a l a r g e nozzle, t h e
r a t e s of temperature change a r e s u f f i c i e n t l y ' small t o permit slow condensation (fog formation) on t h e s e n u c l e i , a p o s s i b i l i t y which i s
suggested i n r e f e r e n c e 5 . The 54' F supercooling i s t h e r e f o r e never
even approached, so t h a t t h e cataclysmic process of formation and
growth of molecular n u c l e i which i s responsible f o r t h e condensation
shock cannot occur.
The slow condensation process produces only a small e f f e c t on t h e
Mach number d i s t r i b u t i o n s , and t h i s e f f e c t i s d i s t r i b u t e d r a t h e r t h a n
concentrated a s i n t h e case of t h e condensation shock. I n order t o
i n v e s t i g a t e t h i s e f f e c t , t h e nozzle was operated with varying humidity
conditions. The corresponding Mach number d i s t r i b u t i o n s a t t h e wall
and near t h e c e n t e r of t h e nozzle a r e shown i n f i g u r e 22, where t h e
r e l a t i v e humidity i n t h e low-speed s e c t i o n of t h e t u n n e l i s increasing
from runs 1 t o 4. The Mach number a t which t h e estimated s a t u r a t i o n
temperature i s reached i s i n d i c a t e d f o r each case. The Mach number i s
seen t o decrease with increasing condensation throughout t h e nozzle.
16
NACA RM L5OA03a
This r e s u l t was a t f i r s t surprising because heat addition, due t o the
condensation, increases the Mach number i n subsonic flow but has t h e
opposite e f f e c t i n supersonic flow. It must be remembered, however,
t h a t the amount of condensation i s not uniform throughout the nozzle
but i s increasing with increasing Mach number (decreasing temperature).
Tne amount of condensation i s therefore g r e a t e r a t the t h r o a t than i n
any subsonic p a r t of the nozzle; but the Mach number a t the t h r o a t
cannot exceed unity. The mass flow i s therefore reduced throughout
the subsonic p a r t of t h e nozzle. The amount of the reduction i n Mach
number a t any s t a t i o n i n the subsonic p a r t due t o the reduction i n
mass flow exceeds the increase due t o condensation a t t h a t s t a t i o n
because t h e condensation a t the t h r o a t i s always g r e a t e r . I n t h e
supersonic region, t h e Mach numbers a r e reduced because the amount
of condensation i s everywhere g r e a t e r than a t the t h r o a t .
Another e f f e c t between runs 1 and 4 ( f i g . 22) i s the movement of
t h e e f f e c t i v e t h r o a t approximately 7 inches downstream. Such an e f f e c t
might be due t o a time l a g i n condensation, but it i s believed due t o
t h e influence of heat t r a n s f e r on the boundary-layer development.
Because of t h e decreasing stagnation temperature between runs 1 and 4,
t h e r e l a t i v e l y increasing heat t r a n s f e r i n t o t h e tunnel increases t h e
r a t e of growth of t h e boundary-layer displacement thickness. I n the
very s e n s i t i v e region near t h e t h r o a t , t h i s e f f e c t may be s u f f i c i e n t
t o s h i f t t h e e f f e c t i v e t h r o a t the observed distance downstream. Such a
s h i f t requires l e s s than 0.00014 increase i n t h e slope of the boundaryl a y e r displacement thickness.
The flow near the center of t h e nozzle i s influenced by condensat i o n i n t h e cooler region near t h e w a l l . This e f f e c t i s evident i n
f i g u r e 23, which shows the dependence of t h e Mach number i n t h e t e s t
s e c t i o n on t h e temperature r e l a t i v e t o s a t u r a t i o n temperature. Note
t h a t the supercooling a t t h e wall i s much g r e a t e r than t h a t a t t h e
center. I n s p i t e of the f a c t t h a t t h e design Mach number on t h e center
l i n e i s t h e same a t t h e 90-inch s t a t i o n (B" i n f i g . 3) a s a t t h e
60-inch s t a t i o n (A i n f i g . 3 ) the decrement due t o condensation i s much
g r e a t e r a t the 90-inch s t a t i o n ( f i g . 23). Thi.s e f f e c t i s due t o t h e
progressive condensation i n t h e region near t h e walls, f o r the flow i n
t h i s region a f f e c t s t h a t downstream along t h e Mach l i n e s .
Wall Interference f o r a Typical
Transonic-Airplane Model
I n order t o obtain an estimate of t h e maximum fuselage length, f o r
a t y p i c a l transonic-airplane model, which could be t e s t e d f r e e of wall
interference e f f e c t s i n the Mach number 1 . 2 nozzle, calculations were
NACA RM L5OA03a
made of the flow p a t t e r n about a simple body of revolution located a t
t h e a x i s of symmetry i n t h e supersonic t e s t section. This body cons i s t e d of a cone of revolution, whose angle (35' included angle) was
i d e n t i c a l with t h e nose angle of t h e model fuselage, followed by a
c y l i n d r i c a l afterbody whose diameter (approx. 3.75 i n . ) was equal t o
t h e maximum diameter of the model. The flow calculations depended,
f o r t h e most p a r t , on the method of c h a r a c t e r i s t i c s but the use of
some approximate theory was a l s o necessary because of t h e existence of
subsonic flow near t h e cone. The r e s u l t s of t h e calculations a r e shown
i n f i g u r e 24. Tests of t h e transonic-airplane model yielded t h e nozzlewall Mach number d i s t r i b u t i o n shown i n f i g u r e 24, which indicated a
measured shock s l i g h t l y upstream of t h e calculated nose-shock l o c a t i o n
a t t h e w a l l . The measured shock appeared a s a gradual compression
because of the t h i c k boundary l a y e r a t t h e nozzle wall. The i n t e r s e c t i o n of t h e r e f l e c t e d shock and t h e body could not be calculated because
of the occurrence of subsonic flow behind t h e shock i n t h e region of
i n t e r s e c t i o n , but t h e maximum. fuselage length was estimated t o be about
one tunnel radius. Consideration of t h e i n t e r f e r e n c e problem by making
use of t h e free-stream Mach l i n e s y i e l d s nonconservative r e s u l t s a s i s
shown i n f i g u r e 24. No measurements were available t o determine t h e
a x i a l l o c a t i o n of t h e r e f l e c t e d shock a t t h e center of t h e tunnel
downstream of t h e model fuselage.
E f f e c t s of Flow Nonunifomities on Model Forces
I n order t o obtain p r a c t i c a l information concerning t h e e f f e c t s
of flow disturbances on model force measurements, a complete model of
a transonic airplane was investigated i n t h e supersonic t e s t s e c t i o n
when t h e flow disturbance a t the 75-inch s t a t i o n on t h e a x i a l center
l i n e produced a deviation of a s much a s 0.05 from the t e s t - s e c t i o n Mach
number of 1.2 ( f i g . 19). The a x i a l l o c a t i o n of t h i s moderately strong
flow disturbance was varied with respect t o t h e model by changing t h e
a x i a l l o c a t i o n of t h e model along t h e center l i n e of t h e t e s t section;
and model f o r c e s were measured with t h e flow disturbance located i n
several regions near the wing and t a i l surfaces. The r e s u l t s of t h i s
investigation, reported separately i n reference 6, indicated t h a t a t
Mach number 1 . 2 no s i g n i f i c a n t changes of l i f t , drag, and pitchingmoment c o e f f i c i e n t s f o r t h e given model were produced by flow nonuniformities equivalent t o Mach number decrements of t h e order of 0.05
or l e s s a t t h e a x i s of symmetry and extending over an a x i a l distance
of 2 o r 3 inches o r 6 t o 10 percent of t h e model length. Reference 6
a l s o contains experimental d a t a which indicate t h a t t h e f l u c t u a t i n g
normal shock i n t h e stream a t the downstream end of t h e supersonic t e s t
s e c t i o n has no s i g n i f i c a n t e f f e c t on model f o r c e s u n t i l it approaches
t h e region of t h e model base and t a i l surfaces.
NACA RM L3OA03a
Power Requirements
The power required t o operate t h e nozzle i s shown i n f i g u r e 25.
Additional power d a t a a r e included from investigations i n t h e diffusing
portion of a l i n e r o r i g i n a l l y i n s t a l l e d i n the tunnel t o f a c i l i t a t e
t e s t i n g at high subsonic speeds and from t h e preliminary rough wooden
nozzle t e s t s . The apparent s c a t t e r i s due t o t h e wide v a r i a t i o n i n
conditions under which the data were obtained. For t h e p l a s t e r nozzle
operating a t Mach number 1 . 2 the power given, about 12,500 horsepower,
i s t h a t required f o r maintenance of the supersonic flow i n t h e t e s t
section. This value was obtained i n winter; with the higher operating
temperature required f o r avoiding condensation e f f e c t s i n summer, t h e
power required f o r the same Mach number would be somewhat g r e a t e r . I n
general, the Mach number values above 1 . 2 were maintained over only
short distances, m d the power values may therefore be somewhat l e s s
than would be required f o r t h e production of a t e s t region a t t h e same
Mach numbers. I n t h e case of t h e o r i g i n a l tunnel with subsonic l i n e r
and a l s o f o r the rough wooden nozzles, shock waves e x i s t i n g i n t h e flow
upstream of t h e terminal normal shock may have a f f e c t e d the power
required. Because t h e Mach number i s i n most cases not constant over
t h e cross section of t h e tunnel, t h e Mach number corresponding t o any
given power value may a l s o depend on the p o s i t i o n a t which t h e Mach
number i s obtained. The power absorbed by t h e strut-and-survey tube
was estimated not t o exceed 850 horsepower.
The estimated power required because of t h e normal shock terminating the supersonic flow i s shown i n f i g u r e 25 f o r comparison. The
shock power a t Mach number 1 . 2 i s s t i l l small and i s only a minor p a r t
of t h e t o t a l power. With increasing Mach number, t h e shoclr-power curve
and t h e course of t h e experimental power values tend toward convergence.
This behavior i s t o be expected i f t h e d i f f u s e r flow i s not spoiled by
t h e shock, because with increasing Mach number the d i f f u s i o n through
t h e normal shock exceeds by an increasing amount the preceding expansion
i n the nozzle, so t h a t decreasing d i f f u s i o n i s required of t h e d i f f u s e r .
A s pointed out i n reference 3, the flow d i d not separate behind t h e
normal shock terminating the supersonic flow with Mach number 1 . 2 .
The e q e r i m e n t a l power data may be compared with nozzle-empty
estimates made without consideration of difyerences i n tunnel geometry
and Reynolds number from information presented i n references 7 t o 9 .
The powir required t o operate t h e 8-foot high-speed tunnel was cons i d e r s b l y l e s s than esttm8ted from the f i ~ s two
t
of these references
( f i g . 25). The upper curve i n f i g u r e 25 was computed f o r a d i a b a t i c
compression from blower pressure r a t i o s given i n reference 7; a tunnelf a n e f f i c i e n c y of 80 percent was assumed. The curve from reference 8
was obtained ,for t e s t - s e c t i o n pressures representative oS those i n
t h e 8-foot high-speed tunnel during operation. Reference 9, which
became available a f t e r completion of t h i s investigation, l e a d s t o a
19
NACA RM L5OA03a
power estimate i n good agreement with t h e present experimental d a t a
( f i g . 2 3 ) . The power estimates contained i n reference 9 were based on
d i f f u s e r - e f f i c i e n c y t e s t r e s u l t s from a number of wind tunnels
including a high-speed subsonic tunnel l a r g e r than t h e 8-foot highspeed tunnel; i n reference 9 c e r t a i n reasonable assumptions were made
regarding t h e v a r i a t i o n of t o t a l - p r e s s u r e l o s s e s with Mach number.
C ONCLUSI9NS
1. A l a r g e , approximately 8-foot-diameter, supersonic nozzle of
c i r c u l a r cross section f o r Mach number 1 . 2 was made t o produce a t e s t
region of s a t i s f a c t o r i l y uniform flow.
2. Considerable care, including compensation f o r boundary-layer
development, was required i n the design and construction, p a r t i c u l a r l y
i n t h e region of t h e t h r o a t .
3. Significant improvement of t h e nozzle flow was achieved by
reducing i r r e g u l a r surface waviness by amounts of the order of
0.006 inch i n height.
4. Experimental Mach number d i s t r i b u t i o n s i n t h e nozzle were i n
excellent agreement with t h e design d i s t r i b u t i o n s . The flow i n t h e
supersonic t e s t section was f r e e of shocks and ot' a degree of
uniformity s a t i s f a c t o r y f o r aerodynamic t e s t i n g ; deviations from the
design Mach number a t the center l i n e d i d not exceed 0.02.
5 . The design f e a t u r e s leading t o the supersonic t e s t s e c t i o n were
consistent with t h e production i n t h e t h r o a t region of a subsonic t e s t
section of s a t i s f a c t o r i l y uniform flow f o r Mach numbers a s great a s 0.99.
The length of the uniform t e s t region was such a t any Mach number t h a t
with fineness r a t i o about 6 t h e s i z e of t h e model t h a t could be t e s t e d
was l i m i t e d by choking r a t h e r than by t h e length of t h e t e s t s e c t i o n .
6. Flow i r r e g u l a r i t i e s , which were propagated along Mach l i n e s ,
tended t o become concentrated near the a x i s of symmetry.
7. Small surface i r r e g u l a r i t i e s of t h e nature of roughness, cracks
i n t h e p l a s t e r surfaces, and small l o c a l i z e d d i s c o n t i n u i t i e s i n slope
produced no noticeable e f f e c t on t h e flow and were therefore concluded
t o be masked by t h e t h i c k boundary l a y e r . An axisymmetrical 0.030-inchdiameter bump around the w a l l i n t h e divergent portion of t h e nozzle
produced a l a r g e disturbance corresponding t o a Mach number deviation
of 0.13 a t t h e a x i s .
NACA RM L5OA03a
8. I n t h i s l a r g e nozzle, no evidence was found of t h e e x i s t e n c e
of a condensation shock. Slow condensation occurred and produced a
d i s t r i b u t e d e f f e c t on t h e Mach numbers.
9. The power r e q u i r e d f o r supersonic o p e r a t i o n was i n s u b s t a n t i a l
agreement with t h e smallest of s e v e r a l e s t i m a t e s obtaiiled from p r e v i o u s l y published information.
10. Except f o r t h e d i r e c t e f f e c t s of d e n s i t y changes, t h e t u r b u l e n t
boundary l a y e r i n t r a n s o n i c flow behaved i n e s s e n t i a l l y t h e sane manner
a s f o r incompressible flow; i t s e f f e c t on t h e o u t s i d e flow w a s equival e n t t o a displacement of t h e streamlines near t h e w a l l by a n amount
equal t o t h e boundary-layer displacement t h i c k n e s s .
Langley Aeronautical Laboratory
National Advisory Committee f o r Aeronautics
Langley Air Force Base, V a .
NACA RM L5OA03a
1. Ferri, Antonio: Elements of Aerodynamics of Supersonic Flows.
Ch. XIII. The Macmillan Co., 1949.
2. Tetervin, Neal: Approximate Formulas for the Computation of
Turbulent Boundary-Layer Momentum Thicknesses in Compressible
,
Flows. NACA ACR ~ 6 ~ 2 21946.
3. Tulin, Marshall P., and Wright, Ray H.: Investigation of Some
Turbulent-Boundary-Layer Velocity Profiles at a Tunnel Wall with
Mach Numbers up to 1.2. NACA RM L9H29a, 1949.
4. Oswatitsch,
Kl. : Formation of Fog in Wind Tunnels and Its Influence
on the Investigation of Models. Reps. and Translations No. 459,
British M.A.P. vzlkenrode, June 1946.
5. Oswatitsch,
Kl. : The Formation of Fog in High-Speed and Supersonic
Tunnels. Reps. and Translations No. 948, British M.A.P.
~Slkenrode,Sept. 1, 1947.
6. Ritchie, Virgil S. : Effects of Certain Flow Nonuniformities on Lift,
Drag, and Pitching Moment for a Transonic-Airplane Model
Investigated at a Mach Number of 1.2 in a Nozzle of Circular
Cross Section. NACA RM L9E20a, 1949.
7. Ackeret , J. :
High-Speed Wind Tunnels. NACA TM 808, 1936.
8. Crocco, Luigi: High-Speed Wind Tunnels. Translation No. 366,
Materiel Command, U. S. Army Air Corps, July 30, 1943.
9. Goethert, Bernhard: Power Requirements ttnd Maximum Reynolds Number
in Transonic Wind Tunnels Having a Maximum Mach Number of 2.0.
5564, Air Materiel Command, Army Air Forces, March 25,
AAF No.
1947
a~allof tunnel d i f f u s e r .
34
36
38
40
42
44
46
48
50
52
54
32
30
24
26
28
20
22
4
6
8
10
12
14
16
18
2
o
Distance downstream
of e f f e c t i v e
minimum section, x
(in. )
46.2500
46.2543
46.2574
46.2602
46.2620
46.2631
46.2669
46.2741
46.2870
46.3054
46.3280
46.3540
46.3827
46.4142
46.4470
46.4810
46.5157
46.5507
46.5855
46.6203
46.6532
46.6860
46.71.69
46.7470
46.7798
46.8043
46.8302
46.8530
radius, r
(in. )
Geometric
Effective
radius, r '
(in. )
Distance downstream
of e f f e c t i v e
minimum section, x
(in. )
COORDINATES FOR DIVERGENT PORTION OF TEE MACH NUMBER 1.2 N0ZZI;E
Geometric
radius, r
(in. )
0
u
l
t+
12
2
&*
P,
W
\%AT
%
--
-------------------------
-------
46.7820
46.6001
46.6157
46.6295
46.6422
46.6532
46.6621.
46.6700
46.6764
46.6819
46.6863
46.6898
46.6922
46.6940
46.6951
46.6960
46.6964
46.6966
46.6967
46.6967
46.6967
46.6967
46.6967
Effective
radius, r '
(in.)
NACA RM L5OA03a
COORDINATES FOR CONTRACTION CONE OR CONVERGENT PORTION O F THI3
MACH NUKBER 1 . 2 NOZZLE
Distance upstream of e f f e c t i v e
minimum section, x
(in. )
Geometric radius, r
(in. )
NACA RM L5OA03 a
CU rn
* G
r-! 0
.ri
$2
P
I!.
G
@
a rn
! iad
a, @
Piri
ld r-!
2 -$rn
@ d m
-P
a,
cd
d
f2@0
G
X N .A
.&-I r-!
0 N
J
NACA RM L 5 O A 0 3 a
'4
v)
Q)
c
.y
3
I
3..
c G#Ik
F
6.
3
Q
B
'h l
\
0
3
X
0
7
I
'
Q
b
l k
$
3
I
\
I
I
I
o
12
1 0
I .?
1 %
I e
Ib
-
Figure 4. Comparison of calculated and measured boundary-layer displacement thicknesses a t the tunnel
wall with subsonic l i n e r i n s t a l l e d i n the t h r o a t region. Mach number i n l i n e r t e s t section, 0.85.
NACA RM L5OA03a
I. 3
1.2
I .I
1.0
'2
.9
-.c
.8
Q
9
F3
'7
.6
-50
/I
----------------
/ --
/
*5
.4
/'
/
.3
/ 7
.2
/
/
/
/
//
.I
/
I
osoo
I
400
--I
//
300
200
/oo
0
I
I
/OO
A x i a l distance from e f f e c t i v e minimum section, x , in.
-
Figure 5.- Calculated d i s t r i b u t i o n of Mach number a x i a l l y along t h e
e n t i r e l e n g t h of t h e c o n t r a c t i o n cone and nozzle.
I
200
NACA RM L5OA03a.
NACA RM L5OA03a
NACA RM L50A03a
NACA RM L 5 O A 0 3 a
Figure 9.- The completed temporary plaster nozzle for Mach number 1.2, as
viewed from downstream.
"///////////////h
I . . .
I.............
I
....
. ..... .....' .... . .... , ..
, ,\
Figure 1 0 . - Arra.ngement f o r measuring t o t a l and s t a t i c p r e s s u r e s near t h e a x i a l c e n t e r l i n e and s t a t i c
p r e s s u r e s a t t h e w a l l of t h e nozzle.
"
/
,.........
Stafic -pf essure orificesin p / w s f e r wall
Sf u f i c - p r e s s u r e o r i f i c e s
in c y l i n d r i c q l f u b e
Figure 11.- Cone
(3O
included angle) f o r measuring t o t a l and s t a t i c pressures along t h e a x i a l center
l i n e of t h e supersonic t e s t section.
a
z
5..
NACA RM L 5 O A 0 3 a
.
.--
Figure 13.: Mach number distributions obtained from static-pressure measurements (at surface of
cylindrical tube) axially along the nozzle center line for subsonic and supersonic flows.
NACA RM L5OA03a
41
42
NACA RM L 5 0 A 0 3 a
2A
+'8
M al
d A
0 -I=
rl
"jg
Arl
rl k
rl I
cd rl
.rl
X
cd
cd
.rl
-P
FI
m a,
d 4J
0 0
.rl pl
+,
5 a,
2k 5
-P F:
m -I=
.rl 4
a b
83
@
"'g
-5 2
z
8;
rl-P
6 d
-P 0
d v
8
a,
*rl J2
k -P
%%
E aO,
'2 !.rl
!Em ha,
.rl -P
k d
cd a,
F4
8
U
u
NACA RM L50A03a
NACA RM L50AQ3a
.gure 18.-Mach number distributions obtained from pressure measurements at approximately 1, 7,
and 13 inches off the axial center line of the supersonic test section with a slight dfsturbance
in the flow.
Much line f o r
Experimen f
Cone A
Cone B
0 Cy//'ndr;ca/ fube
NACA RM L 5 O A 0 3 a
NACA RM L 5 O A 0 3 a
2
-
,
in.
Figure 22.- Introduction of humidity effects in the Mach number distribution by insufficient heating of
the flow mixture.
x
f ~ u m i d i f yeffecfa neqliqible)
Dis fame from design effective minimum sec +ion,
"
U
xI
d---
WU//
A f.
60
90
onches downsfream o f effec five minimum section)
Ax.ia/ locafion
Figure 23.- Effect of supercooling on indicated Mach numbers a t the wall and near the center l i n e of the
supersonic t e s t section.
Difference bef ween safupafion f emperclf ure and femperafure
C Y ~s f a f i o n , fsdurdfian-f,
O F
0-
Near
cenfer
Sfafions uf which / W c h number meusuremenfs were obfained
z
Ca
W
E
\jl
2
3-
2
51
NACA RM L50A03a
6'4
8
'
h ' /k'zb'&'&
' 3 k j ' & ' 4 ! + ' 4 k 1'sh.' 5 k 1 6 b
DisSonce downstreurn of model nose, x, in.
Figure 24.- Calculated flow p a t t e r n about a cone of r e v o l u t i o n with
c y l i n d r i c a l afterbody and comparison of c a l c u l a t e d and measured
shock l o c a t i o n s a t t h e nozzle wall.
NACA RM LfSOA03a
Subsonic l i n e r
/ , / ,
1
/
/
/
/
/
/
/
1
1
1
/
1
1
1
1
/
/
/
1
/
1
1
/
/
1
1
1
1
,
r
1
~
~
~
1
'
Rough wooden noz z /e
y Subsonic / i n e r
/ /
-
/
I
/
/
/
/
/
/
/
/
"
/
/
/
/
/
1
/
~
/
r
/
r
/
J
x
/
~
~
,
x
,
,
,
,
Mach number /. 2 ~ o z /ze
Experimental
data
from 8-foaf high -speed funne/
Subsonic liner
o Tunnel em f y ; M at center
of f y n n e l
S t r u t m difuser; M 4f center of f u n n e l
R o ~ g hwooden nozzle and subsohic liner
at center o f f u n n e l
o S t r u f in diffuserj
d S f r u t ~ndiTu.ser;
at f ~ / n n e lw a l l
fl
Mach number 1.2 M O Z Z I ~
. M at nozzle wall
M at cen f e r o f nozz le
M 4f n o z z l e w a l l
F i m e 25.- Power r e q u i r e d f o r t r a n s o n i c o p e r a t i o n of t h e Langley 8-foot
high-speed t u n n e l .