Obery, Leonard J

Wise, George A

English

NASA Technical Reports Server

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RM E52I02 
RESEARCH MEMORANDUM 
INFLUENCE OF FUSELAGE-MOUNTED ROCKET BOOSTERS ON FLOW 
FIELD AT INLET AND ON DIFFUSER PERFORMANCE OF 
STR UT-MOUNTED ENGINE AT MACH NUMBERS 
OF 1.8 and 2.0 
By George A. Wise and Leonard J. Obery 
Lewis Flight Propulsion Laboratory 
Cleveland, Ohio 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
October 29, 1952 
Declassified September 17, 1958 
• , 
https://ntrs.nasa.gov/search.jsp?R=19930087265 2020-06-17T11:29:27+00:00Z

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NACA RM E52I02 
NATIONAL ADVISORY COMMITrEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
INFLUENCE OF FUSELAGE-MOUNTED ROCKE'l' BOOSTERS ON FLOW FIELD 
AT INLET AND ON DIFFUSER PERFORMANCE OF STRUT-MOUNTED 
ENGINE AT MACH NUMBERS OF 1.8 AND 2.0 
By George A. Wise and Leonard J. Obery 
SUMMARY 
The effect of fuselage-mounted rocket boosters on a strut-mounted 
engine was investigated in the Lewis 8- by 6-foot supersonic wind 
tunnel at Mach numbers of 1 .8 and 2.0 and a Reynolds number of approx-
imately 48X106 , based on body length. The boosters were pairs of 
uircular cylinders with conical fore bodies mounted on the top and the 
bottom of the fuselage, and the engine was mouhted on a horizontal 
strut. For the investigation, the boosters were located in two 
longitudinal pOSitions and fairings were added in the forward 
position. 
The results of the investigation indicated that the boosters in 
the forward position had the most adverse effect on engine performance. 
Either moving the boosters aft or adding fairings was effective in 
reducing the losses in engine mass flow and pressure recovery, but 
the fairings were more effective. 
INTRODUCTION 
An auxiliary power system is required for boosting a ram-jet 
missile to some operating condition, at which point the ram jets 
furnish the necessary thrust. In many cases, rockets are used as 
this auxiliary power system; for supersonic missiles, these rockets 
may become quite large relative to the missile size. The mounting of 
these large bodies on the fuselage of a missile may affect the engine 
performance. 
Therefore, an investigation was conducted to determine some of 
the interference effects on engine performance resulting from rocket 
boosters mounted on the fuselage. Pairs of rocket boosters were 
mounted on the top and bottom of a fuselage together with a nacelle 
engine strut-mounted on the side of the fuselage. The purpose of 
the investigation was to determine the effect of the boosters on the 
flow field at the inlet station and the extent to which the engine 
performance was affe cted. 
2 NACA RM E52I 02 
The investiga tion was conducted in the Lewi s 8- by 6- foot super-
sonic wind tunnel a t Mach numbers of 1. 8 and 2.0 through a range of angl es 
of atta ck and mass flow ratios. The Reynolds number of the investi gati on 
was approxi ma tely 48Xl06 based on body length . 
APPARAWS AND PROCEDURE 
The model investi gated i n the tunnel (fig . 1) cons i sted of a body 
of revoluti on with pa irs of dummy rocket boosters mounted on the top and 
bottom of the fuselage and a nacelle - type engi ne strut- mounted hori zon-
tall y on the body . The symmetri cal fuselage had a length- di ameter rati o 
of 12 and a maxi mum diameter of 9 inches . The dummy r ocket boosters had 
600 coni cal noses and cyli ndri cal afterbodi es 4 . 7 i nches i n di ameter 
and wer e of arbi trary length . Each pair of boosters was connected by a 
meta l pla te a cross the top of the cones . They were located at two longi-
tudi na l stati ons as shown i n f i gure 1 (hereinafter called boosters - forwar d 
and boosters- aft locations ) and, with the boosters in the forward pos i -
tion, a f a iri ng (fig . 2 (a)) was placed over the nose of the boosters . 
Photographs showing the fai ring on the booster s and the boosters i n the 
aft pos i t i on are presented in f i gur e 3 . 
The engi ne was located It engine diameters from the body center line 
and was mounted in the horizonta l plane (fi gs . 1 and 3 ). The dif f user 
was identi cal to the modified di f f user of r eference 1 with the excepti on 
of the i nner body aft of the cyli ndri cal portion of the outer shell . 
Coordi na tes for the di ffus er and the engine di mensions are gi ven i n 
figur e 2 (b ). The mass flow t hr ough the engi ne was controlled by a mova -
ble plug mounted from the rear of the body . 
A f l ow survey rake was mounted on the opposi te s i de of the body f rom 
the engi ne and was located longi tudi nally in the plane of the i nl et . 
Fi gure 2 ( c ) shows the details of the flow survey i nstr umenta t i on . Pitot 
press ur e tubes wer e mounted ad j a cent to 60 flow survey wedges . The 
enti re sur vey apparatus was shi f ted ver tical l y to pr ovide a f l ow s ur vey 
over the area shown i n f i gure l ( c ) . I nstrumenta t i on for the engi nes con-
s i sted of sta t i c pr ess ur e r akes loca ted at the di f fu ser exi t and in the 
combusti on chamb er, designa ted i n f i gure 2(b ) as s t a t i ons 3 and A, 
r especti vely . 
I n the reducti on of t he da t a f rom the flow sur vey appa r atus , the 
measured pi tot pressur es wer e corre cted for normal shock loss es by mea ns 
of the local Mach nUlnber s a s measured by the wedges . As can be seen 
from figur e 1 (fr ont vi ew of t he model ), shi fting the sur vey apparatus 
vertica lly permi tted the local Mach number s and t otal press ures to be 
mea sur ed at the srune points in t he f l ow f ield. Si dewash angl e s wer e mea s -
ured di rectly with the wedges a nd these data wer e corrected for wedge 
misa l i nement or possible f ree - s t r eam angula r i ty by subtracti ng the 
NACA RM E52I02 3 
sidewash for t he body alone at zero angle of attack from the rest of the 
data. The mass flow through the engines was computed from the known 
exit area and the measured combustion-chamber static pressure, assuming 
choking at t he exit area. The ratio of the exit area to the eombustion-
chamber area determined the combustion-chamber Mach number, which 
together ,nth the measured static pressure determined the total pressure 
at station A. This total pressure was assumed to act at the exit station. 
Total pressure recovery for the diffuser was determined from t he known 
mass flow and the measured static pressure at the diffuser exit (sta-
tion 3) . 
RESULTS AND DISCUSSION 
The charac t eristics of the flow field at a Mach number of 2.0 are 
presented in figures 4 to 6. With the boosters located in the for"ward 
position (fig . 4(a))) the inlet was immersed in a region of decreased 
available total pressure. The addition of the fairings to t he boosters 
in the forward position considerably reduced the loss in available total 
pressure and also confined it to a region cl ose to the body (fig . 4(b)). 
Moving t he boosters aft also reduced the loss in available total pres-
sure although the effects of the boosters on the flow field extended 
nearly to the inlet (fig . 4(c)). 
For the boosters forward, an outwash of the order of 20 was meas-
ured in the vicinity of t he inlet (fig. 4(a)). Addition of the fairings 
completely eliminated the outwash a t the inlet and produced a small 
amount of inwash as indicated in figure 4(b). Moving the boosters rear-
ward also eliminated the outwash produced by the boosters-forward con-
figuration; however, the inwash resulting from the boosters-aft configu-
ration was greater in magnitude t han the previously measured outwash 
(fig. 4 (c)). Thus, moving the boosters rearward decreased the total 
pressure loss, but at the expense of increased sidewash. 
For each configuration, it can be seen that a variation of angle of 
attack has only slight effect on the flow field (figs. 4 to 6 ). This is 
evidenced by the fact that the regions of low total pressure and large 
sidewash are) in general, located in the same position with respect to 
t he inlet throughout the angl e - of-attack range. Because of this small 
angle-of-attack effect, the changes in configuration had the same effect 
at angles of attack of 30 and 60 as at 00 . 
For a Mach number of 1 .8 and an angle of attack of 00 (fig. 7), the 
lowest available total pressure occurred with the boosters in the forward 
position. Also, as in t~e case of a Mach number of 2.0) the engines were 
well outside the low total pressure region when the fairings were on the 
boosters. 
4 NACA RM E52I02 
The sidewash contours show that the regions of outflow moved out-
board of the body from their positions a t a Mach number of 2 .0. At a 
Mach number of 2 .0 with the boosters aft, t he engine inlet was immersed 
almost entirely in inflow, whereas at a Mach number of 1.8 a large por-
tion of the inlet was immersed i n outflow . The other configurations 
also showed the same trend . Because of this shifting, the grea test side-
wash at the inlet occurred with the boosters in the forward pos i tion . 
Also, greater sidewash was noted with the fair i ngs on at Mach number 1.8 
than at Mach number 2 .0. 
A comparison of engine mas s flow and pressure recovery for various 
configurations a t zer o angle of attack for the two Mach numbers is pre-
sented in fi gure 8 . At a Mach number of 2 .0, the boosters in the aft 
position and the forward position with fairings had a small but meas-
urable effect on mass flow and pressure recovery . The boosters in the 
forward position, however , reduced the pressure recovery and the maxi -
mum mass flow ratio and completely eliminated the stable range . 
At a Mach number of 1 .8, the boosters in the aft posi t ion caused a 
relatively greater decrease in mass flow and pressure recovery than at 
a Mach number of 2 .0. This is due probably to the fact that the shock 
off the boosters l ay f arther forward a t a Mach number of 1 . 8 than at a 
Mach number of 2 .0. Also, with the boosters forward, there was a stable 
operating range a t a Mach number of 1 .8 where there was none at a Mach 
number of 2 .0. 
The engine characteristics for all configurations at a Mach number 
of 2.0 are presented in figure 9 . With the boosters off, the pressure 
recoveries and mass flows were nearly t he same as those obtained with the 
engine al one (reference 1) except that the stable operating range was 
reduced at angle of attack. 
Mounting the boosters on the fuselage in the forward position reduced 
both t he maximum total pressure recovery and the mass flow ratio of the 
diffuser . Also, at 00 and 30 angles of attack, the stable subcritical 
r ange obtained with the boosters- off configuration was entirely eliminated. 
The l a ck of stable subcrit ical range (as contrasted to the stable range 
obtained for other configurations at these angles of attack) is believed 
to result from the region of low avail able total pressure located i n a 
restricted section near the cowl l i p, a s shown in the flow surveys of 
figures 4( a) and 5(a) , Since, when the low total pressure regi on receded 
from the lip at 60 angle of attack, a small stable subcritical range was 
obtained . Comparison of the sidewash contours for the angle- of- attack 
range and among the t hree configurations i ndicat es that sidewash probably 
did not cause t he diffuser instability . With t he boosters forward, increas-
ing the angle of atta ck from 30 to 60 increased t he mass flow ratio and p r o-
vided a limited s table subcritical r a nge but decrea sed the total pr es sure 
recovery. Considering the flow sur vey (f igs. 5(a) and 6(a) ), it appears 
NACA RM E52I02 5 
possible that the increased mass flow ratio resulted from an increase in 
available total pressure and a decrease in sidewash. Although not pre-
sented, the Mach number in the vicinity of the inlet was also lower for 
an angle of attack of 60 than for 00 and 30 • It is also possible that the 
increase in mass flow could have resulted from downwash changes. 
With the fairings on the boosters, the only characteristic that 
varied with angle of attack was the stable operating range. This vari-
ation, which was a decrease of stable subcritical range with an increase 
of angle of attack, was apparently not the result of changes in avail-
able total pressure or sidewash but is possibly a function of downwash. 
With t he boosters in the aft position little change was noted in the 
mass fl ow or pressure recovery as the angle of attack was increased from 
00 to 30 . A considerable decrease, however, resulted from an increase 
in the angle of attack from 30 to 60 . Apparently this variation in dif-
fuser characteristics with angle of attack cannot be attributed to either 
the available total pressure or the sidewash, since, from figures 4(c), 
5(c), and 6 (c)) little change is noted in these quantities through the 
angle-of-attack range. I t can also be seen that for angles of attack 
of 00 and 60 the mass flow and pressure recovery were reduced approxi-
mately 2 percent from the 00 and 60 values for the boosters-off configu-
ration. It thus may be concluded that some other stream variable, such 
as downwash or Mach number, favorably influenced the diffuser at an angle 
of attack of 30 b~t not at 00 or 60 • 
SUMMARY OF RESULTS 
An investigation to deteI111ine the effects of fuselage-mounted rocket 
boosters on the flow field at the inlet and on the diffuser performance 
of a strut-mounted engine at Mach numbers of 1.8 and 2.0 was conducted 
in t he Lewis 8- by 6-foot supersonic wind tunnel. The boosters, which 
were pairs of circular cylinders with conical forebodies, were located 
on the top and bottom of the fuselage. They were i nvestigated in two 
longitudinal positions and with fairings. 
The following results were obtained: 
1. At a Mach number of 2 . 0, the boosters in the forward position 
had the effect of immersing the inlet in a region of low available total 
pressure. Placing f airings on the boosters or moving them aft without 
fairings tended to reduce the loss in available total pressure. However, 
moving the boosters aft increased the available total pressure at the 
expense of increased sidewash. At a Mach number of 1.8, the greatest 
loss in available total pressure and the largest sidewash were obtained 
with the boosters forward. Addi ng 1'airings to t he boosters was the most 
successful way of reducing these loss es . 
6 NACA RM E52I02 
2. At both Mach numbers the greatest losses in mass flow and total 
pressure recovery of the engine were obtained with the boosters in the 
forward position. The most effective means of reducing these adverse 
effects was the placing of fairin~s over the nose of the boosters. 
Moving the boosters aft was also helpful but not so effective as the 
fairings. 
3. Angle of attack had only a slight effect on the flow field 
characteristics for each configuration. 
Lewis Flight Propulsion Laboratory 
National Advisory Committee for Aeronautics 
Cleveland, Ohio 
REFERENCE 
1. Obery, L. J., and Krasnow, H. S.: Influence of a Canard-Type Control 
Surface on the Internal and External Performance Characteristics of 
Nacelle-Mounted Supersonic Diffusers (Conical Centerbody) at a 
Rearward Body Station for a Mach Number of 2.0. NACA RM E52F16 , 1952. 
NACA RM E52I02 
Top 
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Side 
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Support strut 
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, Boosters-forward ~ ! ~ i station 55 .2":------, l I "---Engine - inlet station 74.1" 
Station L Boosters - aft s tat ion 61.2" 
o 
Front view at 
station 74.1 
Engine 
1" 14"4 ------I 
Area of survey 
r-'''-1 ~ 
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F igure 1 . - Sketch of model showi ng l ocation of var ious components . 
7 
8 NACA RM E52I02 
t 
4 . 7 " 
! 
(a ) Fair ing . ~ CO · 2785 
x , i n. 0 0 . 15 0 .40 0 . 65 1.15 1. 65 2 . 15 2 . 3 4 3 . 15 1 4. 15 1 6 . 90 1 7 . 65 1 9 . 65 11 . 06 
R in. 1. 50 1. 54 1. 60 1. 66 1. 74 1. 79 1. 84 1. 85 ~Straight t aper to ) 2 . 48 
r , i n. 0 . 86 0 . 93 1. 03 1.13 1. 26 1. 33 1. 38 1.40 1.46 11. 511 1.56 11. 54 11. 43 1.3 1 
6 . 365 
O.D. 
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(b ) Engine . 
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Figure 2 . Det a i ls of fai r ing , engine, and s urvey r ake . 
.. _. _ ------
(a) Boosters t"OIvard with fairings . ~~ 
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Figure 3 . - Photographs of model mounted in tunnel. 
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Figure 3 . - Concluded. Photographs of model mounted in tunnel. 
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NACA RM E52I02 
Total pressure 
ratio 
(percent of f r ee stream) 
90 
(outwash is positive) 
(a) Boosters forward . 
(b) Boosters with fairings . 
( c ) Boosters aft. 
Figure 4 . - Contours of total pressure ratio and sidewash for three configurations . 
Angle of attack, 00 ; Mach number, 2 .0. 
11 
.... 
12 
Total pressure 
ratio 
(percent of stream) 
90 
ACA RM E52I02 
(outwash is positivs) 
(a) Boosters f orward. 
8t-
08S 
------=!-. --e ~ -------==t- . . - 7"'9----t-~ 
(b ) Boosters with fai r ings . 
(c ) Boosters aft . 
Fi gure 5 . - Contours of t otal pressure ratio and sidewash for three configurations. 
Angle of attack, 30 ; Mach number, 2.0 . 
1 
NACA RM E52I02 
Total pressure 
ratio 
(percent of free stream) 
--=l-~e 
la) Boosters forvard. 
(b) Boosters with fairings . 
(e l Boosters aft . 
(outwash is positive) 
2 
1 
Figure S. - Contours of total pressure ratio and sidewash for three configurations. 
Angle of attack, So; Mach number, 2 .0 . 
13 
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14 
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85 
9 
95 
(a ) Boosters forward . 
(b) Boosters with fairings . 
( e ) Boosters aft . 
NACA R~'I E52I02 
(outwash is positive; 
Sidewash 
(deg ) 
Figure 7 . - Contours of total pressure r atio and sidewash fo r three configurations . 
Angle of attack, OO j Mach number, 1 . 8 . 
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NACA RM E52I02 
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Mass flo., ratio, ~/mO 
(b) Mach number, 1 . 8 . 
Figure 8. - Variation of total pressure recovery with mass 
flow ratio f o r four configurations at two Mach numbers . 
Angle of attack, 0°. 
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Mass flow ratio, ~/mo 
(c) Boosters with fairings. (d) Boosters aft . 
Figure 9 . - Variation of total pressure recovery with mass flow ratio for four configurations at several 
angles of attack . Mach number, 2.0. 
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Influence of Fuselage-mounted Rocket Boosters on Flow Field at Inlet and on Diffuser Performance of Strut-mounted Engine at Mach Numbers of 1.8 and 2.0

Influence of Fuselage-mounted Rocket Boosters on Flow Field at Inlet and on Diffuser Performance of Strut-mounted Engine at Mach Numbers of 1.8 and 2.0

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