Mossman, Emmet A

Pfyl, Frank A

English

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RM A55Ll6 
RESEARCH MEMORANDUM 
A STUDY OF A SYMMETRICAL, CIRCULAR, 
INTERNAL COMPRESSION INLET 
By Emmet A. Mossman and Frank A. Pfyl 
Ames Aeronautical Laboratory 
Moffett Field, Calif . 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
February 13J 1956 
Declassified May 29. 1959 
https://ntrs.nasa.gov/search.jsp?R=19930089021 2020-06-17T08:20:38+00:00Z

NACA RM A55L16 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
A STUDY OF A SYMMETRICAL, CIRCULAR, 
INTERNAL COMPRESSION INLET 
By Emmet A. Mossman and Frank A. Pfyl 
SUMMARY 
A preliminary experimental study of symmetrical, Circular, internal 
compression inlets has shown that they attain pressure recovery equal 
to that measured by conical nose inlets at Mach numbers up to about 2.3. 
Tnis pressure recovery was obtained with configurations having essen-
tially zero pressure drag of the external surfaces. 
lNI'RODWTION 
Recently, an experimental investigation has been made at Mach num-
bers up to 2.5 of an inlet which shows promise. The purpose of this 
paper is to present an interim report describing the development of the 
inlet and the progress that has been made to date. 
SYMBOLS 
M Mach number 
m mass flow, lb-sec/ft 
total pressure, lb/sq ft 
A area, sq ft 
p static pressure, lb/sq ft 
x longitudinal distance, ft 
D diameter, ft 
2 NACA RM A55L16 
Subscripts: 
00 free stream 
1 inlet 
max maximum 
min minimum 
c compressor 
t total 
0 entrance 
DISCUSSION 
Up to the present time, external compression inlets have produced 
relatively efficient supersonic com~ression of the induction air, but 
the wave drag of the external cowls has been high . On a typical air-
plane this inlet drag is from 10 percent to 20 ~ercent of the total air-
plane drag at Mach numbers above 2.0. This cowl drag is mainly a function 
of the initial lip angle. It has been found experimentally that the best 
overall performance of conical inlets occurs when the li~ internal sur-
faces are nearly alined with the flow direction immediately behind the 
conical shock wave (fig. 1). This figure shows also the flow angularity 
behind a 300 cone and the angle of shock detachment. If 30 is added for 
lip thickness, the external lip angle a~proaches even closer to the angle 
of shock detachment . These considerations indicate that the drag of 
external compression inlets will be large and will probably increase 
with Mach number because the li~ angles must increase. 
If the lip angles could be kept low, the pressure drag could be 
markedly reduced . The internal compression inlet shown in figure 2 is 
designed for use with a typical jet engine. The resulting external sur-
faces have very low angularity, approximately 10 ; consequently, the wave 
drag would be negligible. The relative dimensions of the nacelle shown 
at the bottom of the figure are for a M = 2.0 design . As the design 
Mach number increases, both the inlet and exit diameters increase relative 
to the engine envelope diameter, and the nacelle will approach even closer 
to a straight tube . A symmetrical circular configuration was selected 
which allows the minimum area of the internal duct to be varied by trans-
lation of the center body. The photographs at the top of f i gure 2 show 
the cone extended for starting, an off-design or transitional pOSition, 
and the design position with the cone fully retracted. The movement 
NACA RM A55L16 3 
of the center body should be programmed with the flight Mach number. 
In addition, an automatic control is necessary which can sense the 
position of the terminal shock wave and actuate the center body to 
maintain the shock pos ition near the minimum area. Control mechanisms 
for supersonic inlets which might be adapted to this internal com-
pression inlet are described in reference 1. The angularities of the 
compression surfaces are low - 80 to 100 for the cone and 1.50 to 2.00 
for the lip annulus . These angles are kept low to avoid shock-induced 
separat i on during the internal compression process by limiting the pres-
sure rise in an incident wave to less than the value which haE been 
found to cause separation . It w0111d be desirable to reduce the length 
of the internal ducting. However, in the present design it has not been 
found necessary to include long stabilizing sections rearward of the 
minimum- area station. Consequently, the present internal compression 
inlet is slightly shorter than equivalent conical inlets. The internal 
compression inlet has been tested only at 00 angle of attack. 
Use of the internal compression inlet will result in a net gain 
only if the pressure recovery is sufficiently high. The conical inlet 
will be used arbitrarily in the following discussion as a standard for 
comparison. The maximum pressure recovery of conical inlets is shown 
in figure 3 for the Mach number range from 1 . 8 to 2 . 5. A solid-line 
curve representing the present state of the art is shown in the figure 
and will be used for comparison purposes. The experimental pressure 
recovery of the internal compression inlet as a function of Mach number 
for four internal shapes is shown in figure 4. Comparison of the data 
with the best external conical-shock inlets shows that the pressure 
recovery is about the same over a range of Mach numbers to 2.3. For 
three of the inlets the compression surfaces were generated by straight 
lines, the ratio of the minimum area to the inlet area being varied. 
The fourth case, which gave the highest pressure recovery at Mach num-
bers greater than 2.1, has a curved center body and a curved lip annulus. 
It should be noted that the contraction ratio for this inlet corresponds 
to that for inlets with straight internal elements which gave low pres-
sure recovery for t he same Mach number range. 
Some idea of why the inlet with the curved center body and curved 
lip annulus gave higher pressure recovery than the other internal com-
pression inlets was gained by mapping the internal flow field by using 
the method of characteristics. Figure 5 shows the shock-wave pattern 
and the computed pressure distribution on the center body and lip annulus 
for one of the inlets with straight internal elements. The shock waves 
from the center body and annulus initially had about the same strength. 
However, the pressure rise at the first reflection of the shock wave from 
the annulus on the center body was much larger than it was for the first 
reflection of the center-body shock on the annulus. The pressure gradi-
ents behind the shock intersections also were dissimilar. In an effort 
to equalize both the pressure ratio across each shock-boundary intersection 
4 NACA RM A55L16 
and the pressure gradients behind the intersections, curved compression 
elements were employed; and as noted previously, the modification was 
moderately successful . It is thought that further improvements are 
possible, especially at higher Mach numbers, by shaping the internal 
surfaces to follow contours derived by the characteristics method. 
Of importance to the inlet designer is the off-design performance. 
On the left -hand side of figure 6 the take-off performance of the 
internal compreSSion inlet is shown. The take-off characteristics are 
similar to those for the conventional conical inlet. The curves on 
the right -hand side of this figure illustrate the "matching" character-
istics of the inlet and engine combination. The ordinate of this curve 
is the ratio of the capture (or entrance) area to the streamtube area 
supplied by the inlet or required by the engine. The air handling 
qualities of the inlet are shown by the solid-line curve. The internal 
compression inlet is assumed to operate as a normal-shock inlet at Mach 
numbers up to 1.6. Above a Mach number of about 1.8, the inlet operates 
with the streamtube area equal to the inlet area and has no spillage 
drag. 
The engine air requirements (in terms of the streamtube- area ratio) 
are shown by the dashed-line curve in figure 6. For this particular 
engine at M above 1.2, a maximum of 3 percent of the inlet air woUld 
have to be bypassed for maximum efficiency. Below a Mach number of 1.2, 
either the rotational speed of the engine could be reduced slightly or 
some loss in 'pressure recovery would be incurred. Many other matching 
programs could be devised, but figure 6 indicates that the problems 
should be no more severe with the internal compression inlet than they 
are with other types of supersonic inlet-engine combinations. 
CONCLUDING REMARKS 
A preliminary experimental study of symmetrical, circular, internal 
compreSSion inlets has shown that they attain pressure recovery equal to 
that measured by conical nose inlets at Mach numbers up to about 2.3. 
This pressure recovery was obtained with configurations having essen-
tially zero pressure drag of the external surfaces. 
Ames Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Moffett Field, Calif., Nov. 2, 1955 
REFERENCE 
1. Wilcox, Fr ed, and Perchonok, Eugene: Aer odynamic Control of .Super-
s onic Inlets f or Optimum Performance . NACA RM E55L14, 1950 . 
.. 
NACA RM A55L16 
LIP ANGLES OF CONICAL INLETS 
40 
0 32 
w 
Cl 
w'24 
-.J 
o 
z 
<l 16 
a.. 
-.J 
8 
~FLOW ANGLE FOR 
30° CONE 
o I 
1.8 2.0 2.2 2.4 2.6 2.8 3.0 3.2 
Moo 
Figure 1 
INTERNAL COMPRESSION INLET 
STARTING OFF DESIGN DESIGN 
r ° "en , 0,,,,,, ° ""'"'E: Ei?f±:~1 
TRANSLATING CONE J TURBINE~ j 
COMPRESSOR NOZZLE 
Figure 2 
5 
6 NACA RM A55L16 
PRESSURE RECOVERY OF CONICAL INLETS 
(ml/moo= 1.0) 
1.00 
.90 
-if. 
+ 
+ 
.80 
Pt/Ptoo 
.70 
+ 
.60 
.50 I I o L LI __ LI __ .LI _----::-'-1 :---::-,1-:----::-,=1 ----::'"::------:::-' 
1.8 1.9 2.0 2.1 2.2 2.3 2.4 2.5 
Moo 
Figure 3 
PRESSURE RECOVERY OF I NTERNAL COMPRESSION INLET 
1.00 DESIGN ACMIN)lAoo 
CONICAL INLETS - 0 .649 
COMPARISON CURVE 0 .539 
.90 <> .476 
x D. { .604 
<t D. CURVED 
:::E SURFACES 
----s .80 
cL 0 
-....... 
u 
£--
.70 ~ 0 0 
.60 
.50 L ~I __ L-_~I __ ~I __ ~I ___ I~_~I_~I 
o 1.8 1.9 2.0 2.1 2.2 2.3 2.4 2.5 
Figure 4 
NACA RM A55L16 
THEORETICAL PRESSURE DISTRIBUTION 
ON COMPRESSION SURFACES 
SHOCK 
PATTERN 
PLOCAL 
Pro 
2.0 
1.0 
CENTERBODY 
/ 
J=..:-=-=-=-=-=-=.ry AN NU LUS 
L' ____ L' ____ ~' ____ ~, ____ ~,L_ __ ~, 
o .4 .8 1.2 1.6 2.0 
X / Do 
Figure 5 
PERFORMANCE OF THE INTERNAL COMPRESSION INLET 
AT OFF-DESIGN CONDITIO NS 
SEA LEVEL 
1.0 2.0 
......... ~/--=€=3 .......... w 
'" 
.9 ~ \ ~ 1.6 
""" 
I ::10 !---k 1 <t w 
Pt
c 
• I :: 1.2 
.8 ~ C/) <{ PteD ...... 
-===: ~ w .8 0:: :::> 
.7 I-0.. 
<t 
.4 u 
<{ 
.6 0
1 LI I I 
o .2 .3 .4 .5 .6 
MENTRANCE 
Figure 6 
NACA - La ngley F ie ld, Va. 
35,000' 
SUPPLIED BY 
INLET (THEORETICAL) 
REQUIRED BY 
TYPICAL ENGINE 
I I I 
1.0 1.4 1.8 
Moo 
I 
2.2 
7 


A Study of a Symmetrical, Circular, Internal Compression Inlet

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