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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

"

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 rut­tank 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 inboard­forward 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 drag­coefficient 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-drag­coefficient 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 inboard­forward 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 con­figurations 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 Super­sonic 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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CON FJ DENTIAL

NACA 65-009 airfoi l normal to L.E.

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Figure 1 .- General arr angement of drag research vehicle with wing tanks . All dimensions in inches.

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TANK L:)CATlm MODEL X Y

R.WING ~? 11.949 12.865 0

103 7.063 12.865 1.089 104 7.063 5.329 -1.089 120 111 .949 5.329 0

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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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14 NACA RM L50A19

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Figure 5. - Variation of total-drag coefficient and tip helix angle with Mach number. Wing of 200 square inches, aspect r atio = 2 .7 and sweep­back angle = 340 .

NACA RM L50A19

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Figure 5.- Continued .

NACA RM L50A19

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Figure 6.- Comparison of tank drag coefficients. The coefficients are based on area of two tanks equal ~ to 13.2 square inches. Tank fineness ratio = 7.44. ~

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