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CJASSIFIED DOCUMENT Restriction/Classification Cancelled InformaHon so Classificd may be imparted only to p e r s o n s in the military and naval services of the United States, appropriate CivrLian officers and employees of the Federal Government who have a leatimate Interest therem, and to United States citizens of known loyalty and discretmn who of necessity must he informed thereof.
NATIONAL ADVISORY COM
WASHINGTON
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? . ? a Restriction/Classification Cancelled NACA RM ~ ~ 5 0 ~ 1 6 a e e a e NATIONAL ADVISORY COMMITTEE F O R AERONAUTICS ~e RESEARCH MESIORANDUM f o r the Bureau of Aeronautics, Department of the Navy T H E ZERO-LIFT DRAG O F SEVERAL CONFIGURATIONS O F TAE XAAM-N-2 PILOTLESS AIRCRAFT TED NO. NACA DE332 By James R. Hall and Carl A . Sandahl S U M M A R Y Free-flight t e s t s have been made t o determine the z e r o - l i f t drag Base- of several configurations of the XAAM-N-2 p i l o t l e s s a i r c r a f t .
pressure measurements were a l s o obtained f o r some of the configurations.
The r e s u l t s show t h a t increasing the wing-thickness r a t i o from 4 t o M = 1.3 and 6 percent increased the wing drag by about 100 percent a t Increasing the nose fineness r a t i o by about 30 percent a t M = 1.8.
from 5.00 t o 6.25 reduced the drag coefficient of the wingless models a maximum of about 0.030 (10 percent) a t M = 2.0. A corresponding change i n nose shape f o r the winged models decreased the drag coeffi- cient by about 0.05 i n the Mach number rarige from 1.1 t o 1.4; a t Mach numbers greater than 1.6 no measurable reduction i n drag coefficient was obtained. The drag of the present Sparrow fuselage is l e s s than that of a parabolic fuselage which could contain the same equipment.
INTRODUCTION A t the request of the Bureau of Aeronautics, Department of the Navy, an investigation of some of the aerodynamic characteristics of several configurations of the XAAM-N-2 (sparrow) i s being conducted u t i l i z i n g f r e e - f l i g h t techniques. The f i r s t phase of the investigation was concerned with the determination of the drag a t zero l i f t of several configurations d i f f e r i n g i n nose fineness and wing-thickness r a t i o s .
This phase of the investigation has been completed and the r e s u l t s a r e reported herein. Also included a r e base-pressure measurements obtained f o r some of the configurations tested.
The f l i g h t t e s t s were conducted a t the P i l o t l e s s Aircraft Research Station a t Wallops Island, Va.
Restriction/Classification Cancelled CONFIDENTIAL NACA RM S L ~ O C I ~ ~ S Y M B O L S drag coefficient based on maximum cross-sectional. area of fuselage (0.442 s q f t ) base-pressure coefficient
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base pressure ambient s t a t i c pressure dynamic pressure TEST VER1CI;ES The t e s t vehicles used i n t h i s investigation were 1.125-scale models of the XAAM-N-2 p i l o t l e s s a i r c r a f t . The models were constructed by the Naval Aircraft Factory a t Philadelphia, Pa. The general arrangement of the models i s shown i n figure 1. A photograph of a typical model is shown i n , figure 2.
The fuselages consisted of an ogival nose section, a cylindrical center section, and a boat-tailed a f t e r section and were made of 0.064-inch-thick duralumin skin with r i n g s t i f f e n e r s .
The wings and f i n s were of duralumin and the wings were bolted t o the fuselage center section by a single trunnion leaving a --inch gap between the wings and fuselage. Details of the wing-body intersection a r e given i n figure 3.
I n table I a r e l i s t e d the configurations tested. The ordinates f o r the two nose shapes tested a r e given i n table 11. The over-all length of the fuselages was held constant; the variation i n nose fineness r a t i o was obtained by varying the point of tangency of the ogival nose and the center section. Model 6 was equipped with a nose telemeter antenna, dimensions of which a r e given i n figure 4. The nose fineness r a t i o of t h i s model i s calculated on the b a s i s of nose shape before being modified by i n s t a l l a t i o n of the antenna. A l l t e s t vehicles were polished before launching.
The models were propelled by an ABL Deacon rocket motor which provided a t o t a l impulse of about 19,800 pounds-seconds over a burning period of approximately 3.5 seconds. A ?-inch HVAR booster was employed f o r model 5 i n order t o obtain data a t higher Mach numbers.
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a Photographs of models without and with booster a r e shown i n . figures 5 and 6, respectively.
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TEST M E T H O D S The t e s t vehicles were tracked by Doppler velocimeter t o obtain flight-path velocity and longitudinal acceleration. An SCR 584 radar s e t w a s used t o obtain the f l i g h t paths. By means of standard NACA telemetry, measurements of t o t a l head, base pressure, and longitudinal and normal acceleration were obtained f o r several of the models.
The drag of the models was determined from values of longitudinal deceleration obtained from the Doppler velocimeter during coasting f l i g h t .
These data, i n conjunction with S C R 584 radar flight-path measurements, Doppler velocimeter measurements of flight-path velocity, and radiosonde observations, were used i n the calculation of the total-drag coefficient a s a function of Mach number. For model 6 the drag was a l s o obtained using telemetered values of longitudinal acceleration and t o t a l head.
The base-pressure coefficient - Pj w a s calculated from
(c, = q
telemetered base-pressure measurements and ambient s t a t i c pressure obtained from f l i g h t path and radiosonde measurements. The flight-path velocity was obtained from telemetered total-head measurements and from Doppler velocimeter. The base-pressure pickup was located on the inside of the afterbody between the rocket nozzle and skin a s shown i n figure 7.
The rocket nozzle and pressure pickup were insulated t o eliminate thermal e f f e c t s on the base-pressure pickup. The afterbody was sealed t o prevent i n t e r n a l a i r flow.
Some typical f l i g h t paths, obtained with the SCR 584 radar s e t , a r e shown i n figure 8. The variation of Reynolds number with Mach number f o r the range of a l t i t u d e and climatic conditions encountered during the t e s t s i s given i n figure 9.
Accuracy of Data Drag coefficient.- The random e r r o r s i n the determination. of CD, a s indicated by the s c a t t e r of the data points i n figure 10, a r e small.
The systematic e r r o r s i n CD derived from Doppler velocimeter may be a s large a s f0.040 and k0.0075 a t Mach numbers of 1.0 and 2.0, respec- tively. The systematic e r r o r s i n CD derived from telemeter measurements may be a s large a s 20.080 and 20.020 a t Mach numbers 1.0 and 2.0, respectively.
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-e6 Base-pressure coefficient.- The random e r r o r s i n Cn, as indicated, by the s c a t t e r of the data points i n figure 11, a r e small. A t the lower Mach numbers investigated, these data a r e subject t o rather high a .
: systematic e r r o r s since the quantity pb - po i s of the same order of
* magnitude a s the r e l i a b i l i t y of the telemetered base-pressure measure- ments. The systematic e r r o r s i n Cpb may be a s large as f O . l and 20.01 at Mach numbers of 1.0 and 2.0, respectively.
RESULTS AND DISCUSSION Drag Measurements The variation of z e r o - l i f t drag coefficient with &ch number f o r the configurations tested i s summarized i n figure 12. The condition of zero lift was substantiated by the normal accelerometer. The r e s u l t s f o r the three wingless models agree within the accuracy of the measurements except a t the highest Mach numbers investigated. The drag coefficient was reduced 0.030 (about 10 percent) a t Mach number 2.0 by increasing the nose fineness r a t i o from 5 t o 6.25. A corresponding change i n nose fineness r a t i o f o r the models having wings of 4-percent-thickness r a t i o reduced the drag coefficient by about 0.05 i n the Mach number range from 1.1 t o 1.4. A t the Mach numbers greater than 1.6 the change i n nose fineness r a t i o had no measurable e f f e c t on the drag of winged models. The increase i n drag due t o increasing the wing-thickness r a t i o from 4 t o 6 percent is obtained from the curves f o r models 2 and 3 i n figure 12. The increase i n wing thickness increased the wing drag (taken as the difference between the drag of the winged models and t h a t of wing- l e s s model 6) by about 100 percent a t M = 1.3 and by about 30 percent at M = 1.8. I n making t h i s comparison it i s assumed that the drag of the wingless model is not affected by the presence of the nose antenna.
Base -Pressure Measurements The variation of base-pressure coefficient ,with Mach number f o r a wingless and two winged models i s given i n figure 11. A t the lower supersonic Mach numbers investigated, the presence of the wings tended t o maintain the base-pressure coefficient a t about -0.1. The base- pressure coefficient of the wingless model approackied zero a s the Mach number approached one. A t the higher Mach numbers investigated a l l three configurations tended t o exhibit a value of base-pressure coefficient of -0.09.
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.. Component Drag Coefficients
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e The contributions of the skin-friction, base-pressure, nose-pressure, and residual-drag coefficients t o the t o t a l drag coefficient measured f o r e .
wingless model 6 a r e shown i n figure 13. The residual-drag coefficient : * i s defined here a s the drag remaining a f t e r the f r i c t i o n , base-pressure, and nose-pressure drag coefficients have been subtracted from the t o t a l drag coefficient and consists of the f i n , b o a t - t a i l pressure and f i n - body interference drag coefficients. The skin-friction drag coefficient was calculated using the value of wetted-area skin-friction coefficient obtained from unpublished measurements of the boundary layer on a large- scale f r e e - f l i g h t t e s t vehicle. The base-pressure drag coefficient was calculated from measured values of the base-pressure coefficient obtained with model 6. The nose-pressure drag coefficient was obtained by the method of Laitone presented i n reference 1.
The p o s s i b i l i t y of reducing the drag by changing the present fuselage t o one of parabolic shape has been considered. A parabolic fuselage which could contain the equipment used i n the actual Sparrow missile would have a maximum diameter of about 9 inches instead of 8 inches located a t the 4.0-percent fuselage s t a t i o n and would taper gradually from t h i s point t o a diameter of about 6.5 inches a t the base.
Such a parabolic shape would have approximately the same skin-friction drag coefficient a s the present shape but would have about 20 percent more nose-pressure drag. I n addition, the very shallow slope of the afterbody of such a parabolic fuselage, according t o unpublished work on the e f f e c t of afterbody shape on base pressure, would induce more base suction. O n the basis of these considerations, it may be concluded t h a t the fuselage tested i s a good one from the ~ t a n d p o i n t ~ o f both low drag and ease of manufacture.
CONCLUSIONS The following conclusions a r e based on the t e s t s of 1.125-scale models of the XAAM-N-2 p i l o t l e s s a i r c r a f t : 1. Increasing the wing-thickness r a t i o from 4 t o 6 percent increased the wing drag by about 100 percent a t M = 1.3 and by about 30 percent a t M = 1.8.
2. Increasing the nose fineness r a t i o from 5.00 t o 6.25 decreased the'drag coefficient of the wingless models a maximum of about 0.030 (10 percent) a t M = 2.0.
CONFIDE3 TIAL 6 CONFIDmTIAL NACA RM SLW16a .
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3 . A corresponding change i n nose shape f o r the winged models
. decreased the drag coefficient by about 0.05 i n the Mach number range
0.0 from 1.1 t o 1.4. A t Mach numbers greater than 1.6, no measurable 0. reduction i n drag coefficient was obtained.
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4. The present fuselage has l e s s drag than a parabolic fuselage which could contain the same equipment.
Langley Aeronautical Laboratory National Advisory Committee f o r Aeronautics Langley A i r Force Base, Va.
James R. Hall Aeronautical Research S c i e n t i s t Carl A . Sandahl Aeronautical Aerodynamics S c i e n t i s t Approved : Robert R . Gilruth Chief of P i l o t l e s s Aircraft Research Division REFERENCE 1. Laitone, E. V.: The Linearized Subsonic and Supersonic Flow about Inclined Slender Bodies of Revolution. Jour. Aero. Sci., vol. 14, no. 11, NOV. 1947, pp. 631-642.
NACA RM ~ ~ 5 0 ~ 1 6 a 0.
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SUMMARY OF CONFIGURAmONS TESTED D .
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Wing-thickness Nose fineness Wingle s s 69.0 4 5 .OO 75 5 68.8 6.25 Wingless 75 5 76.8 6 6.25 Wingless 70.3 CONFIDENTIAL CONFIDENTIAL NACA RM S L ~ O C I ~ ~ TABU I1 NOSE COORDINATES if , • •• •• •• •• • •• •
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Figure 2.- A typical XAAM-N-2 model. Nose fineness ratio, 5.00; wing thickness ratio, 0.04.
NACA S L ~ C C I . ~ ~ ...
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• •• •• • •• • ••• • NA CA RM SL 5OC16a C ONFIIENTIAL ••••
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Figure 5.- Unboosted XAAM-N-2 model on launcher.
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Figure 6.- Bo o sted XAAM-N- 2 model on laun c her.
CONFIDENTIAL NACA RM S L ~ O C ~ G ~ m . . .
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for baste pressure pickup
Pressure tap -'\
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Figure 7 . - Sketch of base-pressure pickup i n s t a l l a t i o n .
CONFIDENTIAL NACA RM ~ ~ 5 0 ~ 1 6 a C ONFImNTI.AI, Horizontal distance, ft , Figure 8.- Typical f l i g h t paths obtained with SCR 584 radar s e t .
Numbers r e f e r t o time a f t e r f i r i n g .
Mach number Figure 9.- Variation of Reynolds number with Mach number f o r range of t e s t conditions.
Model fineness I I O . 8 1.0 1 . 2 1 . 4 1 . 6 1 . 8 2 . 0 2 . 2 Mach number (a) W i e d models, velocimeter data points.
Mach number (b) Wingless models, velocimeter data points.
.6 4 Telemeter points 0 D Velocimeter points , U . 4 c.
0 . 2 a 1 . 8 2 . 0 2 . 2 .8 1 . 0 1 . 2 1 . 4 1 . 6 Mach number ( c ) W w l e s s model 6; nose fineness ratio, 6.25.
Figure 10.- Typical drag-coefficient-data points.
Mach Number, M Figure 11.- Variation of base-pressure coefficient with Mach number.
Mach number, M Figure 12.- Variation of drag coefficient with Mach number.
* e e e e i e e NACA RM S L ~ S C C ~ ~ ~ eeo.
m Restriction/ Classification Cancelled Restriction/Classification Cancelled