Morrow, John D

Welsh, Clement J

English

NASA Technical Reports Server

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RESEARCH MEMORANDUM 
EFFECT OF WING-TANK LOCATION ON THE DRAG AND TRIM OF A 
SWEPT-i.VING MODEL AS MEASURED IN FLIGHT 
AT TRANSONIC SPEEDS 
By Clement J. Welsh and John D. Morrow 
Langley Aeronautical Laboratory 
Langley Air Force Base, Va. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
April 4, 1950 
https://ntrs.nasa.gov/search.jsp?R=19930086176 2020-06-17T13:22:56+00:00Z

" 
NACA RM L50Al9 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANIJUM 
EFFECT OF WING-TANK LOCATION ON THE DRAG AND TRIM OF A 
SWEPI'-WING MODEL AS MEASURED IN FLIGHT 
AT TRANSONIC SPEEDS 
By Clement J. Welsh and John D. Morrow 
SUMMARY 
Results of an exploratory f r ee - flight investigation at zero lift 
of several rocket-powe r ed drag r esearch models equipped with wing tanks 
are presented fo r a Mach number range from about 0 .50 to 1.15 . The 
tanks, which wer e slender bodies of r evolution, were mounted on 
340 sweptback, nontaper ed wings of 2 .7 aspect r atio. The tanks wer e 
directl y attached to the wings in such a way t hat their center lines 
wer e positioned on or vertically displac ed from the wing- chord plane 
for tip and inboard spanwise l ocations . The tanks positioned on the 
chord plane wer e also located mor e forward than we r e the ver t ically 
displaced tanks . 
These data show that the test configuration with tanks located 
inboard on the chord line and in the forward position gave the l east 
drag of the four configurations tested . The drag rise for this model 
followed very closely the drag rise of the tankless model. The s t rut-
tank model from a previous paper (NACA RM L8H31a) had a higher drag and 
a drag rise occurring at a l ower Mach number than any of the models 
te sted in this i nvestigation . The r esults of this investigation 
indicate that the tank location has a large effect on the total drag of 
the configuration. The data also indicate that the unsymmetrical models 
experienced a t rim change in the Mach number range from 0 .85 to 1.00. 
INTRODUCTION 
A need exi s t s for experimental data in the transonic speed r egion 
for the prediction of drag char acteristics of general Wing-nacelle and 
external - stores combinations . The Langley Pilotle ss Aircraft Re search 
Division has completed a preliminary program using rocket-power ed 
r esear ch models f r om which t he drag and rate of r oll (a measure of trim 
change ) r e sulting from various tank locations wer e r ecorded. This paper 
contains information obtained from investigations of models having 
2 NACA RM L50Al9 
untapered, 340 sweptback wings of 2 .7 aspect ratio with bodies of 
revolution mounted at differ ent positions on the wings. Configurations 
wer e tested with the tanks l ocated at the tip and inboard, on and 
displaced from the chord plane . The tanks on the chord plane wer e 
l ocated farther forward than the tanks which wer e displaced f r om the 
chord plane . The data are presented as plots of drag coefficient and 
Wing- tip helix angle against Mach number. From these data the drag and 
an indication of t rim changes r esulting from the addition of the tanks 
can be determined. The r esults of this investigation are compared with 
data obtained in a previous investigation which used similar models wi t h 
and without strut -mounted bodies of revoluti on (refer enc e 1). 
The average Reynolds number variation for the models tested in this 
investigation covers a range of from 2 .9 x 106 at a Mach number of 0 .5 
to 8.69 X 106 at a Mach number of 1.20. 
~ 
2V 
p 
b 
v 
M 
A 
S 
R 
SYMBOLS 
wing-tip helix angle , radians 
rolling velocity, radians per second 
total span of 25 .73 inches 
velocity along flight path, feet per second 
total-drag coeffici ent based on exposed wing ar ea of 
200 square inches 
drag coefficient of tanks based on frontal area of two tanks 
of 13.2 square inche s 
Mach number 
b 2 
aspect ratiO, S 
total wing area to center line of body, 248.22 square inches 
Reynolds number based on wing chord of 9.647 inches 
• 
NACA EM L50Al9 3 
MODELS 
The general arrangement of the drag r esearch vehicles used in the 
present investigation is shown in figures 1 and 2 and photographs of 
the models are shown in figure 3 . The basic model construction, 
described in r efer enc e 2, has been altered in the 102 B, 103 A and B, 
and 104 A and B models by the substitution of spinsonde noses, 
reference 3, for the ordinary wooden noses. The tanks were located on 
a 340 sweptback wing with NACA 65-009 airfoil section normal to the 
leading edge. The tanks, of wooden fabrication, wer e of similar design 
t o those used on fighter - type a ircraft and were attached to the wing in 
the relative positions indicated in figures 1 and 2. The tanks had a 
constant fineness ratio of 7.44 and the ratio of tank diameter to body 
diameter was 0 .582 . For convenience, the table in figure 1 shows the 
different tank locations for the fo".IT different arrangements. 
Eight models were used in the investigation. Two models (102 A 
and B) had their tanks located at the wing tips wi th the chord line of 
the wing coinciding with the center line of the tank and with the ends 
of the tanks being flush with the trailing edge of the wing; two other 
• models (120 A and B) had their tanks l ocated at an inboard position with 
the ends of the tanks being flush with the trailing edge of the wing and 
with the center line of the tank coinciding with the chord line of the 
wing (these models will be referred to in this paper as the inboard-
forward symmetrical models); two models (103 A and B) had their tanks 
located at the wing tips with the tanks located on opposite surfaces of 
the wing and with the trailing edge of the tanks extending behind the 
trailing edge of the wing; the final models (104 A and B) had their tanks 
loce,ted at an inboard position with the tanks located on opposite surfaces 
of the wing and with the trailing edge of the tanks extending behind the 
trailing edge of the wing. The tanks, which wer e mounted on opposite 
surfaces, were located in that manner in order that the models would 
maintain straight-line f light paths despite any trim changes that might 
be induced by the tanks and in order to allow a determination of the 
trim- change tendencies by the simple measurement of rolling velOCity. 
The models were propelled by 3 .25-inch aircraft rocket motors which 
were contained within the fuselage . At a preignition temperature of 690 F, 
the rocket motors furnished approximately 2200 pounds of thrust for about 
o .87 second. 
TESTS 
The models wer e flown at the Langley Pilotless Aircraft Research 
Station, Wallops Island, Va. The testing technique whereby drag-
coefficient data are obtained has been adequately described in reference 4. 
4 NACA RM L50Al9 
The accuracy of the drag coefficients is estimated to be ±0.002 at Mach 
numbers above 1.0 and ±0.003 at Mach numbers below 1.0. The accuracy 
of the Mach number is estimated to be within ±O.Ol. • 
angle 
The rolling velocity of each model and the r esulting wing-tip helix 
~ wer e determined by the technique de scribed in r efer ence 3 . 
2V 
The accuracy of the quantity ~ is e stimated to be within ±0.005 
radian throughout the Mach number range . The erratic variation in ~ 
above 0 .9 Mach number in model 102 is not clearly understood. 
The average Reynolds number of the e ight models based on wing chord 
(9.647 inches) parallel to the body center line varied from 2.92 x 106 
at a Mach number of 0.5 up to 8 .69 x 106 at a Mach number of 1.20 . A 
pl ot of Reynolds number against Mach number 1s shown in figure 4 . 
RESULTS AND DISCUSSION 
Drag 
The total -drag coefficient CD and wing-tip helix angle ~ 
ar e presented in figure 5 plotted against Mach number M for the models 
investigated . No drag data were obtained for one of the 104 models nor 
were there any ~ data for e ither of the 120 models. Previous data 
have been obtained for the strut-tank model, a tankless model, and a 
wingless model; these data have been presented in referenc e 1 and are 
included in this paper for comparison . The curves for the strut - tank 
model and tankless model have been slightly modified by use of a later, 
more precise method of r educing flight - test data. The total-drag-
coefficient curve for the wingless model has been included in figure 5 
i n order that the percent of wing-tank-combination drag which could be 
expected to be due to the wing alone may be e stimated . 
The curves shown in figure 5 indicate that the pres ence of the t anks 
caused the drag rise to occur at approximately 0 .03 Mach number lower 
than the drag rise of the tankless model in all of the configurations 
investigated except the inboard-forward symmetrical case . For this 
configuration, the drag rise occurred approximately at the same Mach 
number as the drag rise of t he tankless model . References 5 and 6 may 
partially expla in why the inboard- forward symmetrical model gave the more 
favo r abl e effect of the configurations tested . The r e sult s of those 
, 
NACA RM L50Al9 5 
r efer ences indicated that locating the wing aft of the maximum diame ter 
of a body gave less drag than a forward location. 
An estimated tank-drag- coefficient curve for an isolated tank is 
pres ented in figure 6, which was obtained from a drag curve for a body 
of revolution, reported in reference 7, similar to that of the test 
tanks. The body in referenc e 7 was a fin-stabilized parabolic body of 
revolution with a cut- off stern. Its fineness ratio was 6, and its 
maximum diameter was located at 60 percent of body l ength. The fin and 
base drag was subtracted from the total drag of this body l eaving the 
drag curve shown . It is believed that this curve r epr esents a body 
which is sufficiently similar to the test tanks for comparative purposes. 
The tank-drag coefficient due to the addition of the tanks, which 
included interfer ence effects, was determined by the drag differences 
between the tank-on and tank-off configurations and is also shown in 
figure 6. This coefficient is based on the frontal area of two tanks. 
The variation of the drag-coefficient increment with the different 
models indicate s the importance of tank location with r espect to the 
wing and body in order to minimize tank drag. The inboard-forward 
symmetrical model tanks gave the most favorable drag increment of the 
four configurations investigated. The favorable effects of this tank 
locat ion might not be r ealized if used in conjunction with another type 
body or wing. The drag increments for the models tested were much 
l ower than that obtained from the strut-tank model from reference 1. 
The tanks were located on struts at approximately midspan; however, the 
tank-drag-coefficient curve of the strut-tank model included the drag 
due to the strut. 
Trim Change 
An indication of the trim changes due to the tanks is given by the 
variations of E£ with Mach number presented in figure 5 . The 
2V 
variations of ~ with Mach number for the unsymmetrical models 
indicated that they experienced a trim change in the Mach number range 
from 0.85 to 1 .00 . The r oll obtained at M < 0 .9 for the symmetrically 
located tanks of model 102 is believed due to accidental asymmetries in 
the model; however, the erratic variation in ~ at M > 0.9 for this 
'ZV 
model is not clearly understood. 
6 NACA RM L50Al9 
CONCLUDING REMARKS 
An exploratory rocket - powered flight investigation of drag r e search 
models with wing tanks has oeen conducted near zero lift for a Mach 
number range from 0 .50 to 1.15. The tanks , which wer e slender bodies 
of r evolution, wer e mounted on 340 sweptback, nontapered wings of 2 .7 
aspect ratio in varied positions . The addition of the tanks to the 
models increased the drag coefficient; however, the tanks on the inboard-
forward symmetrical model produced the least increase in drag . Attachffient 
of the tanks also caused the drag rise to occur at 0.03 lower Mach number 
in all models except the inboard-forward symmetrical model . The drag 
rise of this model followed very closely the drag rise of the tankless 
model . The data indicated that the location of the tanks has a marked 
effect on the total drag and also on the point at which the drag rise 
occurs in the Mach numoer range covered in this investigation. Although 
the inboard-forward symmetrical model gave the lowest drag of the con-
figurations tested, it is quite possible that some other tank-wing-body 
combination would ive even lower drag. The data also showed that the 
unsymmetrical models experienced a trim change in the Mach number range 
of 0 .85 to 1.00 . 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Air Force Base, Va . 
NACA RM L50Al9 
REFERENCES 
1. Alexander, Sidney R.: Effect of Strut-Mounted Wing Tanks on the 
Drag of NACA RM-2 Test Vehicle s in Flight at Transonic Speeds. 
NACA RM L8H31a, 1948. 
2. Alexander, Sidney R., and Katz, Ellis: Drag Characteri stics of 
Rectangular and Swept~Back NACA 65-009 Airfoils Havi~g Aspect 
Ratios of 1.5 and 2.7 as Determined by Flight Tes t s at Super-
sonic Speeds. NACA RM L6J16, 1946. 
7 
3. Sandahl, Carl A., and Marino, Alfred A.: Free-Flight Inve stigation 
of Control Effectiveness of Full-Span 0.2-Chord Plain Ailerons 
at High Subsonic, Transonic, and Supersonic Speeds to Det ermine 
Some Effects of Section Thickness and Wing Sweepback. NACA 
RM L7002, 1947. 
4. Katz, Ellis R.: Results of Flight Tests at Supersonic Speeds to 
Determine the Effect of Body Nose Fineness Ratio on Body and 
Wing Drag. NACA RM L 7B19, 1947. 
5. Mathews, Charles W., and Thompson, Jim Rogers: Comparison of the 
Transonic Drag Characteristics of Two Wing-Body Combinations 
Differing Only in the Location of the 450 Sweptback Wing. 
NACA RM L7I01, 1947. 
6 " . Zobel, Theodor: Grundsatzlich neue Wege zur Leistungssteigerung von 
Schneliflugzeilgen . Schr . 1079/44 g., Deutsche Akademie der 
Luftfahrtforschung (Berlin), 1944. (Translation available from 
CADO, Wright-Patterson Air Force Base, ATI9877.) 
7. Katz, Ellis R.: Flight Investigation at High-Subsonic, Transonic, and 
Supersonic Speeds to Determine Zero-Lift Drag of Bodies of 
Revolution Having Finenes s Ratio of 6 .04 and Varyi ng Positions 
of Maximum Diameter . NACA RM L9F02, 1949. 
FRONT VIEW 
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CONFIDENTIAL 
PLAN VIEW 
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Figure 1 .- General arr angement of drag research vehicle with wing tanks . All dimensions in inches. 
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104 7.063 5.329 -1.089 
120 111 .949 5.329 0 
~7 
Figure 2.- Details of wing-tank installation on model. Tank fineness ratio 7.44. Table shows 
different tank locations. All dimensions in inches. 
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chord, 9.647 inches. ~ 
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Mach number. Wing of 200 square inches, aspect r atio = 2 .7 and sweep-
back angle = 340 . 
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Figure 5.- Continued . 
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to 13.2 square inches. Tank fineness ratio = 7.44. ~ 
~ 


Effect of Wing-tank Location on the Drag and Trim of a Swept-wing Model as Measured in Flight at Transonic Speeds

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