Mossman, E. A.

English

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- 
SIJMWRY OF NACA S U l 3 M E Z G ~ - ~  INVESTIGATIONS 
Ames Aeronautical Laboratory 
On msny existing snd proposed airplanes the fuselage shape i s  
aeeuming a g e s t s r  i-qcYrLmce A domj~ailt factor  de teml in ix  the 
shape of the fu~wlage fo-r a y u ~ ~ l i i t - t y p e  turbojet  ail-plm-e ?my be 
the duc t b g  375 ten. gsnn,ral f ~ i x d m m t a l  ~ ~ o q u i r e ~ ~ e n t s  t o  be 
s a t i s f i e 3  3y j e t  airpbmo d~:?Eing ayatem are hi* efiicioncy of the 
impact oii ram preasuro conversion and smell external drag coefficient.  
The impo~tzace of ram recovery can be visualized by c o n s i d e r i ~  i ts 
e f fec t  on a typlcal  9urauit-type airpl.a?e, powered by a jot  engine and . ' 
traveling 6~ miles p3r how a t  aea level .  Analysis shows tha t  f o r  
every 10 percent; decrease i n  ram recovery a t  t h i s  speed the not thrus t  
decreases 7 percent and the specif ic  f u e l  cons tq t ion  increases about 
5 percent. The resul tant  adverse effects on rawe,  climb, and rnaxim~un 
speed are  quite large.  
Recognizing the need for a new type i n l e t  ~ d ~ i c h  would combine the 
good qual i t ies  of the nose i n l e t  w i t h  the short  internal  ducting of 
the external scoop, We National flpviaory C d t t e e  fy Aeronautics 
has developed what is known m a submerged aiy ' in le t  . This intake 
i s  shown in  f i g r e  1 and t h o  component part8 are noted in  t h i s  f igure.  
The entrance i s  ccanpletely sv-Merged below the contovr of the f u s e l a ~ e  
o r  w a l l  in to  which it is  placed. The air t ravels  d-o~rn an inclined 
surface, which Tie have termed the ramp. Roml, a n s o  i a  the angle of 
the interoec tirm of the ram3 f l o o r  w i t h  the fuselage skin, ramp wall 
divergencs is Ene divergence af the ramp side wall from the para l le l ,  
and width-tc-deptii r e t i o  is filmply the r a t i o  of the corresponQing 
dimension..; oF '&e i n i f - t .  This paper sunnnarizes the r e su l t s  of research 
on NACA av%~lerf.;ed inl.zts the 7- by 10-foot, the 40- by 80-foot, and 
the 16-fo(;-t; i r ~ ~ : - ~ ; e s u  tmel  sections a t  t h o  h e s  Aeronautical 
Lab or  a t  o;r . 
A n  entry with para l le l  ramp walls was the f i r s t  t o  be investigated, 
these t e s t s  having been conducted in a m a l l  wlnd channel. A s  expected, 
the pressure recovery with t h i s  parallel-walled entry vas not very ~ o o d ,  
especially a t  the low maas +flow'ratios. It was then reasoned tha t  
ahaping the wall@ to  conform t o  the streamlines a t  some dosired'mass 
flow r a t i o  might r e su l t  i n  be t t e r  duct chmacter i s t ics .  Such on entry 
w i t h  divergent ramp walls was designed and testeb. 
A comparison of the pressure recovery f o r  t h i s  i n l e t  w i t h  tha t  
f o r  a parallel-tralled intake is  shorn i n  f i w e  2. Thsee data were 
obtained from a fu l l - sca le  duct ins ta l la t ion  i n  the, Ames 40- by 
80-foot tunnel; The ordinate f o r  these curves is  ram-recovery r a t io ,  
which is  the r a t i o  of the ram pressure recovered t o  the ram pxeesure 
https://ntrs.nasa.gov/search.jsp?R=19790079925 2020-03-21T20:14:40+00:00Z
available. This r a t i o  was selected because it is re la t ive ly  constant 
for subsonic Mach numbers and 13 readily measurable. The abscissa 
i s  mass flow ra t io ,  which i n  the i n c q r e s s i b l e  case is equivalent to  
i n l e t  velocity ratio.- Conparism of the ram recoveries f o r  the para l le l -  
and divergent-walled Intalses indicates the t  a considerable increase 
r e s u l t s  from the use of divergent w a l l s  a t  the low W S B  flow r a t i o s  
both a t  the entrance.and a t  the compressor. 
. - 
The principal cause of the lower ram recovexy f o r  ea i n l e t  with 
p a r a l l e l  walls, especially in the nass flow range l e s s  than 1.0, is  
the rapid growth of the boundary layer due t o  the adverse pr, a s swe  
gradient along the ramp. Such is not the case f o r  ttze i n l e t  with 
divergent walls. Even though the pressure p a d i e n t  is no l e s s  adverse, 
surveys a t  the entrance sho~i that the boundary layer  on the f loor  of 
the divergentvwalled *let  starts anew and remains re la t ive ly  thin, 
despite the adverse gradient. This probably accounts f o r  the difference 
i n  ram recovery a t  the low mass flow ra t ios  bet~reen divergent and 
parallel-walled submerced i n l e t s ,  The pressure losses wlth divergence 
have a d i f fe rent  origin.  The boundary layer on the f u s e l q e  skin, 
outside the ramp, is  pa r t i a l ly  kept from flo- over the divergent 
ramp edges by two factoro. The f i rs t  of these is the pressure gradient 
over the rear  40 percent of tho remp, the preesures i n  t h i s  region 
being greater than those of the surface in to  which the i n l e t  is placed. 
The second factor is  . that  the ou.tside boundmj layer  does not flow 
over the aharp edge of the ramp wall as  easi ly  as it does with the 
edge rounded. T h e  cause f o r  t h i s  is  not ful1.y understood. 
The pressure losses a t  the entrance f o r  an NACA submerged i n l e t  
a re  concentrated i n  two symmetrical regions, as sho~..m i n  f igure 3 .  
A major p a r t  of these pressure losses agsears t o  oriainnte From a 
turbulent mixlng process s e t  up by a change i n  the flow direction a s  
indicated i n  this same figure. It is probable Chat some of the outside 
boundary layer  is enmeshed and becomes o par t  of mis disturbance. 
An extensive investigation has been made t o  determine the e f fec t  
of modifications on submerged in l e t s .  Varidtions i n  ramp angle, ramp- 
w a l l  divergence, width-to-depth r a t io ,  r q - f X a o r  shape,and boundary- 
layer  thiolrnese have been tested. Results a m  ~ i v e n  in reference 1. 
An evaluation of these data indicates t h a t  sat isfactory duct character- 
i s t i c s  may be obtained f o r  a range of the t e s t  variables. It appears 
t h a t  an optimum design of these i n l e t s  should employ curved diverglrq 
ramp walls, a ramp angle between 5' and 7O, and a width-to-depth r a t i o  
of from 3 t o  5 .  From measured l i p  and ramp pressures, high c r i t i c a l  
speeds were estimated. 
The drag at t r ibutable  t o  t h i e  type of i n l e t  i s  shown In f igure 4 
1 as a functicm of mass flow ra t io .  These data were obtained on .a.;-scale 
J 
typ ica l  duct ins ta l la t ion  on a f igh te r  airpSane. The drag coeff ic ients  
are baeed on wing area. For &n airplane using a 24c j e t  engine and . 
operating a t  a high-speed design mass flow r a t i o  of 0.60, there is  
no incremental change i n  airplane Cd, due t o  t h e  duct instS\llation. ' 
b order t o  check the val idi ty  of the small scalo moasurements, 
models ident ical  except f o r  scale were desi,gied and tosted in both .  
the Ames 7- by 10-foot and. the Amee 40- by 80-foot t~m~ls .  The 
1 ageement between We =.scale and fu l l - sca le  t e s t s  is shown i n  f igure 5. 
2 
The duct location used i n  t h i s  investigation is  noted. i n  the f i g w e  
These data show excellent agreement between-the two t e s t s  and indicate 
t h a t  with proper design hi& ram recoveries a re  at ta inable  on full- 
scale installations'.  (~eynolds number baaed on duct depth .) It might 
be added fur ther  tha t  the variation of ram recovery r a t i o  with angle 
of attack was slidit f o r  these and other ins ta l la t ions  . 
One especially impor ta t  aspect of t h i s  study concerned the 
e f fec ts  of high-speed f l i g h t  on the operation of this type duct. Tests 
of a duct i n ~ t a l l s t i o n  q a &acale model of a f i @ t e r  airplane have 
been made in the h e s  16-foot high-speed tunnel. The r e s h t s  of these 
t e s t e  i l l u s t r a t e  the effect  of Mach number and of the location of the 
i n l e t  on the fuselage. The e f fec t  of Mach number is shown h f igure 6, 
f o r  constant mass flow ra t ios .  The recovery remains essent ial ly  the 
same f o r  the ent i re  Mach. nmber range of the t e s t e  (from 0.3 to  a 
maximum of 0.875). It has not yet  been determine& how high a Mach 
number can be at ta ined while still  maintainirq the high preseure recoverge 
In f igure 7 it mag be seen tha t  a t  a Mach number of 0.875 the c r i t i c a l  
pressure coefficient was juet reached slow the f ron t  of the ramp f loor .  
A shock disturbance would probably f i r a t  occur in t h i s  high velocity 
region when the free-stream Mach number becam somGhat greater 
than 0.875. ~t may be seen that th ia  region extends only over a amall 
portion of the duct width, the flow outside of the ramp on the fuselage 
skin being s t i l l  subsonic. Because the distvrbance takes place over 
a amsuer duct wldth, it mi&t then be indicated tha t  the pressure 
recovery would decrease l e s s  severly f o r  a divergent-walled entry than 
for a parallel-walled one, once the airplane q e e d  wes increased 
enough so tha t  a shock wave f orzaed alcmg tho ranp. 
Given in f igure 8 are  the four i n l e t  locations on the fuselage 
b 
1 of the --scale model, i n  percent of the root  chord forward or  rearward 4 
of the w i n g  l e a d i w  e d ~ e .  It may be seen tha t  the recovery decreases 
s l ight ly  ae the i n l e t  is mved rearward. This was expected since the 
boundary-layer thicknese Xncreaaes i n  t h i s  direction. men  though the 
decrease in ram recovery r a t i o  may not be considered prohibitive, caution 
should be exercised In m o r l q  the i n l e t  reamma.  Primary conaideration 
should be given the flow f i e l d  in to  which the M e t  is  placed. A t  the 
f a r t h e s t  rear position (58.4 percent root chord) the flow a l o w  the 
basic fuselage 'became sonic a t  about the same time as the flow on the 
wing. When We duct was located a t  t h i s  position, the presswe recovery 
began t o  decrease when the Mach number exceeded. ,0.80. This drop becsme 
much more marked a t  moderate angles of attack, We f1.w on the aide 
of the fuselage a b n ~ p t l y  e e p a r a t i x  due apparently t o  the p o s i t i m  and 
intensi ty  of the shock wave on the xing. 
\ 
In conclusion, NACA subnerged i n l e t s  may be designed t o  obtain 
high ram recovery a t  a low resul tan t  drag. Hi@-speed tests on a 
1 6 - ~ ~ a l e  model showed tha t  f o r  t h i s  in s t a l l a t ion  the ram recovery 
remained essent ial ly  constant up to  a Mach number of 0.875. 
I . Moasman, Ermnet A ., and Randall, Lauros M a  : An E~perimental 
Inves t iga t ion  on the Design Variables far NACA Submerged Ehtrancee. 
NACA RM NO . An30, 1947. 
mFigure i.- Schematic view of an NACA submerged inlet installation.
1.00
.90
Hi -PO
HO-PO
(ENTRANCE).80
.70
0
1.00
.90
H2- Po
Ho-po
(COMPRESSOR)
,80
I
/-i- DIVERGENT WA
.L
I / I/---PARALLEL WILLS
J
///
/
/
•2 .4 .6
,2
.e ,.o ,.2 ,.4-
- I t
__ !,VERGELT WALL -
.. __ PARALLEL wALSu
//I
•4 .6 .8 1.0 1.2 1.4
mJm 0
(3_ m -2 •
w/d'4
RAMP ANGLE • 7"
Figure 2.- Comparison of the ram recovery ratio for divergent- and
parallel-walled submerged inlets. Full-scale duct installation.
q
Figure 3 . -  Ram-recovery-ratio contours at the entrance of an NACA 
submerged duct installation. 
P 
CD BASED ON- - 
w . 4  
%AMP A)(OLEI 7. WING AREA= 240 SO. FT DUCT AREA = 1.43 SQ. FT 
Figure 4.- Incremental change in airplane drag coefficient due to an 
9 
NACA submerged duct installation. I- scale model of a fighter 5 
airplane. -
sri 1.1 
'-1 1 7  1 16.75 2 ROOT CHORD 
x 8 0  TESTS 1 7 I I 
. 9  - 
I I 
_I__ 
HI -Po L4bx80  TESTS 
-
no- PO a SO* I 
.a  - "',/m, = 0 . 6 5  
0 .I .2 .3 .4 .S .6 .7 
R x 10'' 
1 Figure 5.- Entrance ram recovery ratio for comparison of --scale 5 
and full-scale NACA submerged duct installations. 
CURVED DIVERGENCE 
w/d = 4 
a = - 2 '  
D U C T  AREA =. I422 SQ. F T. 
F W D .  LOCATION = -16 .7% 
3 .4 .5 . 6  .7 .8 .9 
-MA-CH NUMBER, M 
Figure 6.- Effect of Mach number on the ram recovery ratio at the 
entrance for a +-scale NACA submerged inlet installation 
, J "  . *' 
W i d  = 4.0 n 
RAMP ANGLE = 7' ---- BASIC FUSELAGE 
OC = -2' 
m h o =  0 .53  - . I : .  - DUCT INSTALLED 
M 0.875 
.2 I I 
OUTSIDE OF RAMP 
Figure 7.- Pressure distribution along the ramp of an NACA submerged 
1 inlet installation. ;i. - scale typical fighter airplane. 
MACH NUMBER, M 
Figure 8. - Effect of an NACA submerged duct location on the fuselage 
1 
of a ,- - scale typical fighter airplane. 


Summary of NACA submerged-inlet investigations

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