copy ;:. .: ( RM SL56G23

RESEARCH MEMORANDUM for the

U. S. Air Force

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;DRAG AT MODEL TRIl4 LIFT OF A l/l~-SCAIZ

" ;‘< ": CONVAIR 33-s SUPERSONIC BOMBER

ik iBy Russell N. Hopko and W illiam H. Kinard I :

Langley Aeronautical Laboratory Langley Field, Va.

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lllls material contains frdormatlon affecting tb2 Natfonal Defense of the UnIted stales witidn tb? Of the eqmnage laws. Title 18,u.s.c.,sxs.793 and 764,tka transmissiolrorrevehtionofwhichinaily znanmr to an unauthorized person is prohibited by law.

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON I

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1 .~i b NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS . . ' i 1 < i I

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RESEARCHMEMORANDUM c . . . .'$

for the .-I -3. t r U. S. Air Force .', 'L..

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DRAG AT MODEL TRIM LIFT OF A l/13-SCALE .' ' ,; ',.'

CONVAIR ~-58 SUPERSONIC BOMBER ' I.. :' ,,d .'_ L ..,' Y

By Russell N. Hopko and Will iam H. Kinard .: .!.-- ..' LT. ..,..!

SUMMARY i . : \Q

1 :,’ ,’ . 4, An investigation has been made'by the Langley Pilotless Aircraftq : ":Research Division utilizing a l/13-scale rocket-propelled model of the ,: Convair ~-58 supersonic bomber. The drag at model trim lift was obtained

at Mach numbers between 0.85 and 2.0 at corresponding Reynolds number per foot of 3.5 x 106 and 13.7'~ 106,,respectively. The results of the pres- ent investigation are compared with unpublished data obtained.from several facilities, WADC lo-foot tunnel, Ames 6- by 6-foot supersonic tunnel and

I! the Iangley 16-foot transonic tunnel. A comparison of the drag at tran- sonic speeds and at approximately the same Reynolds numbers showed excel- lent agreement. A drag coefficient of 0.028 at a Mach number of 2.0 was obtained at zero-lift conditions.

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INTRODUCTION /I .": i i, .'

At the request of the U. S. Air Force, the'langley Pilotless Aircraft Research Division has undertaken B flight test program to determine the drag near zero lift of the Convsir ~-58 composite'airplane. The vehicle portion'of the B-58 supersonic bomb& consists of two parts;, the basic inhabited airframe.which is designated the return component, and an expendable pod which is an air-to-surface missile. The complete vehicle with pod attached on a short pylon is designated the composite airplane.

The return component consists of a 60~ modified delta wing at 3O of incidence incorporating 0.13-local-semispan leading-edge camber, a Mach number 2.0 supersonic uarea rule" fuselage, and a swept tapered vertical tail mounted atop the fuselage aft of the wing trailing edge. Four

i' nacelles are pylon mounted underneath the wing.

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The expendable pod is essential ly a body-of-revolution m issile with the support ing pylon attached to its upper surface. Aerodynamic surfaces of the pod, canard and wing, are 60' deltas with-100 swept trailing edges and the'vertical tails are swept and tapered.

Results of some previous investigations during the development of the B-58 have been reported in references 1 to.3. In the present inves- tigation ti l /15-scale rocket-propelled composite model was flown and data were obtained over a Mach number range of 0.85 to 2.0 at corresponding Reynolds number per foot of 3.5 X lo6 and 13.7 X 106, respectively. This investigation was conducted at the Langley Pilotless Aircraft Research Station at Gallops Island, Va.

A

:'A1 .,

I3

cc

CD

CD1

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cm

b

CN

M

2

SYMBOLS,

cross-sectional area

longitudinal acceleration, ft/sec2

acceleration due to gravity,,ft/sec2

chord-force coefficient, Chord force qs

Drag drag coefficient, - qs

internal drag coefficient, Internal drag @

external drag coefficient,, C% - CDI - C!DB

base drag coefficient, po-pD,x&' q% ' s

total drag coefficient, C!~ +,CDI + (3%'

span

normal-force coefficient, Normal force ss

Mach number

overall length

-

NACA RM sL56G23 3

r;

'b

static pressure, lb/sq ft

base pressure, lb/sq ft

9 dynamic pressure, lb/sq ft

R Reynolds number

S total wing area including body intercept, 6.86 sq ft

%I area of nacelle base (four nacelles).

t time .,

V velocity, ft/sec

W ' model weight, lb

7 angle between instantaneous flight path and the horizontal, deg

,

MODEL ',

The general arrangement of the model is shown in figure 1 and a photograph of the model is shown in figure 2: Other pertinent physical characteristics are.presented in tables I, II, and III. The wing, con- structed mainly of steel, has a diamond plan form with 60~ sweep of the leading edge and a -loo sweep of the trailing edge. Outboard of sta- tion 3.33 the wing has an NACA ood+.o8-63 airfoil section; at the root it has an NACA 0~3.46-64.069 section. The wing has 30 of incidence and dihedral of 2°13'4511 outboard of station 3.767; 'The,camber has been designed for a lift coefficient of 0.22 at a Mach number of 1.414. The elevon deflection was O":

The pod wing and canard are of similar plan form to the wing of the return~cxxnponent and have NACA 0004.5-64 airfoil sections. These sur- faces were at 0' deflection for the present investigation.

The vertical tail is a swept tapered surface with an NACA 0005-64 air- foil section. The leading-edge sweep is 52O and the taper ratio is 0.324,

The pOa tail is a swept tapered surface with an NACA 0005-64 section. The leading edge is swept 600 and the taper ratio is 0.35..

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The model was constructed by the -dated-Vu%ee-A4~ Ft. Worth, Texas.

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4 NACA RM SL56G23

TEST PROCEDURE

Instrumentation

The models were internally instrumented by the Langley Aeronautical Laboratory of the National Advisory Conrmittee for Aeronautics with &n eight-channel telemeter which transmitted the following information: longitudinal acceleration (two instruments),, normal acceleration, trans- verse acceleration, total pressure (two instruments), statid pressure, and base pressure. 'The base pressure measurements were made on one inboard nacelle by using four pressure orifices manifolded together and , connected to a,pressure pickup instrument; A modified SC&584 radar unit w&s used'to determine the space'position of the,model in flight. The velocity was obtained with a CW Doppler velocimeter and a rawinsonde

..provided atmospheric conditions and winds aloft velocities throughout .the altitude range transversed by the model in flight.

Propulsion

The model attained a maximum Mach number of approximately 2.0 with an M-5 Jato (Nike booster)., After burnout the booster drag separated from the model and data were obtained during coasting flight.: A photograph of the model in launching position is shown in figure 3.

DATA REDUCTION

Ground Radar

Drag coefficients were obtained during model flight by evaluating the following expression

CD.‘=

V is the velocity velocimeter and corr rected to the tangentialvelocity along the flight path and also car-'L)

cted for winds of the altitud:s.in flight. /.-.,/---e-.e-.4- -- 1" i"-/ i/-..------- - 4.

Telemeter

The longitudinal accelerometer data were used in the following equation

AL W/s %=g q

v

z

a 1 p /:I !S p: j I I “1 1;

NACA RM ~~56~23 5

A similar expression was usea to evaluate the normal and side-force coefficients using the normal a& transverse accelerations, respectively.

The base drag coefficients were determined from

%s= PO - pb s,

q, s

ACCUBACY

The accuracies in coefficient form fdr the Mach number, &rag, and normal-force,data are estimated to be

REKWLTS AND DISCUSSION

The Reynolds numbers per foot are given in figure 4. The total drag coefficients for the configuration are shown in figure 5. These drag coefficients were determined by both CW Doppler radar and telem- etered acce&erations. The data range' of the Doppler radar was from M= 2to M= 1.5 for this investigation; the data range of the telem- eter was from M = 2 to M = 0.85. Excellent agreement was obtained between the Doppler and telemeter data. Base pressures were measured on one inboard nacelle and these data redud&d to drag coefficient are shown in figure 6. Also shown in figure 6 are base, pressure measurements on both'the inboard and outboard nacelles, obtained in t&e Langley 16-foot iransonic tunnel and the Ames 6- by 6-foot supersonic tdnnel. Inasmuch as the outboard nacelle base pressures were not measured in f;light and because the comparison of the flight' and tunnel base pressure measure- ments for the inboard nacelle shows excellent agreement, the outboard nacelle base pressure measurements obtained iiz the Langley 16-foot tran- sonic tunnel were employed in evaluating the external drag data for the present rocket model.

No measurements of the internal drag were made with the ,flight model. The internal drag measurements obtained in the Langley 16-foot transonic tunnel on both the inboard and outboqd nacelles are shoti in figure 7. Also shown.are internal drag measurements obtained in the Langley 4- by k-foot supersonic pressure tunnel on one,outboard nacelle of' a similar

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6 NACA RM SL56G23 .

nacelle configuration. The inlet spike for the 16-foot-transonic-tunnel investigation was set at the M = 0.9 cruise position giving a mass- flow ratio of approximately 0.9. For the present investigation the inlet spikes were set for approximately the flight condition at M = 2 with a respective mass-flow ratio of 1.0 and the mass-flow ratios of subsonic speeds were approximately 0.7. Therefore, calcuiated values of the inter- nal drag using one-dimensional-flow theory and also reference 4 were used to determine the external drag in this investigation.

A comparison of available external drag data is made in figure 8tih~4sh~tis J&e drag coefficient at model trim lift (shown in fig. 9) and essentially

. zero sideslip, as determined from the transverse acceleSations$%s pre- sented and compared with data obtained at WADC lo-foot tunnel, Langley l6-foot transonic tunnel, and Ames 6- by 6-foot supersonic tunnel; the a&a of the latter configuration were obtained with the ,inboard nacelles parallel to the wing chord and the outboard nacelles at -5O to the,wing chord. The configuration tested in the Langley l6-foot transonic tunnel Was, similar to the configuration of this investigation. A comparison of the external drag obtained in this investigation with the results, obtained in the l6-foot trapsonic tunnel at approximately the same Reynolds numbers showed excellent agreement. The 6- by 6-foot tests were m&e with fixed transition ana 9 AC, of approximately 0.0014 due to the boundary-layer trip was estimated. This increment.in drag coefficient has not been subtracted from the data obtained in the Ames 6- by 6-foot supersonic tunnel presented herein. .'

A nondimensional cross-sectional area diagram of the present con- figuration is shown. in figure 10.

CONCLUDING REMARKS

The drag at &Id trim lift of the Convair ~-58 supersonic bomber was obtained'at Mach numbers between 0.85 and 2.0 at corresponding Reynolds number per foot of'3.5 x 106 and 13.7 x 106. The external drag. of the model at trim lift has been compared with data obtained in the hngley l6-foot transonic tunnel. A comparison of the drag at transonic speeds &a at approximately the same Reynolds number showed excellent agreement. A drag coefficient of 0.028 at a Mach number of 2.0 w&s

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obtained at .zero-lift conditions. The model had a mild transonic trim change,~Oka--d-~~-ri-se-Maeh-n~be-r~f~pF~~t~~

Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Langley Field, Va., June 27, 1956.

Aeronautical Reslarch Scielitist

Aeronautical Re'kearch Scientist

Division at3

REFERENCES

1. Hall, James R., and Hopko, Russell N.: Drag and Static Stability at Low Lift of Rocket-Powered Models of the Convair ~~-1626 Airplane at Mach Numbers From 0.7 to 1.5. NACA RM SL53FOga, U. S. Air Force, 1953.

2. Swihart, John M., and Foss, Willard E., Jr.: Transonic Aerodynamic and Trim Characteristics of a Multi-Engine Delta-Wing Airplane Model. NACA RM L5512'i'b, 1956.

3. Hopko, Russell N.: Drag,,Near Zero Lift of a l/T-Scale Moael of the Convair B-58 External Store ati Measured in Free Flight Between Mach Numbers of 0.8 and 2.45. NACA RM SL55G22a, U. S. Air Force, 1955.

4. Fraenkel, 'L. E.: Some Curves for Use in Calculations of the Perform- ance of Conical Centrebbdy Intakes s$ Supersonic Speeds and at Full Mass Flow. Tech. Note No. Aero. 2135, British R.A.E., Dec. 1931.

TARIE I

WING GNCMEl’RY

trailing-edge radius, 0.016 typical; 6ee figure l(b)]

3 4 I 2 5

T T -I- l- l- span station 3.767 chord - 36.216

span station io.667 Chord = 23.048

Span station 15.835 chord = 13.189

spt~~ station 18.150 Chord = 8.768

s&I station 20.667 chord - 3.965

Root chord Chord = 42.741

&per rdinatc

-0.589 -.547 ::g

1:::; -.35? -.320 -.256 -.2c% -.171 -.I39 -al7 -.O% -A35 -.I230 -.oa

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imeIlSiOI! Lower ~rdinatt

Upper rdinste

Lover rdinatf

Upper rclimte

Lower rdimtr

Lawer rdlnatl

Upper rdlnatr

Lower cdlnatc

mP= rtinate

-0.304 -.080

A61 J79 .293 .355 ,421

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0

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kYlinat istance

0 .534

1.06g 2.137 3.206 ;.;g+

8:4&3 12.822 1-f.3a 21.009 24.631 28.233 31.874 35.460

E$2 42:741

0 A16 .@89

1.81.1 2.716 3.6~ 5.432

1%; 1414% 18.103 a.729

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-0.063 434

:Eg

$2 .667 .7u .739

222 .580 .46g .335 .l82

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::g

::g -.335 -.182 -.W 0

0 .288 .576

1.152 1.728 2.304 3.461 4.6u 6.915 9.219

Il.525 13.830 16.134 18.438 20.744 a.896 23.048

.E. radius

-0.304

::E ~381 -.392 :::g -.451

::$ -.4a -.368 -.296 -.w -.u5 -.I%4 0

0 ~65 -.277 .559

1.333 2.064 2.712

a:$

7:913 9.232

10.551 U.187 12.529 13.189

.E. radius

-0.424 -.301 -.253 -.I59

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2-g .256

:Z .2ll .171

:E .036

0

.024

-0.452

-::;: -.466 -.425 -.382

::g; -.262 -.242 -.2ll -.171 -2.22 gz66 -.036

0 ,189

,973 1.333 2.005.

L.E. radius

-0.509

::g -.235 -.163 -.lOl -.013

:E .125

:g .024

0 .016

-0.509

::z -A88 -459 ::;g

::;g -.w -.@a -.051 -.024 0

0 .051 .099 .197 .299 ,397 .595 .792

:%T 1:9% 2.379 2.776 3.173 3.568 3.765 3.965

S.E. radius

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-.592

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1:;::

:::;z -.363 1::; -.21g -.173 -.l28 -.107 -.og1

0 .224 .301 .401 .468

1% .648 .715 ,739

2%

2:; .335 .182 .O%

0

TABIEII

FflJBELAcE GEcMGLllY

rie flgure l(lA -I

Base Ine a ltatio io.occ !4.33?

0 ,835

1.32: 1.30: 1.M 1.67t 1.81b

:%z 2il.5

::tZ 2.81

::g; 1.7l.z 1.52l 1.255

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Ba3e ine 8‘ tatio 3.333 7.666

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tatio 7.80~ 2.133

3ase tne at ;atio1 5.667 L.000

3 A48

:$ .828 .975

1.086 !.173 L.233 L.317 1.353 1.353 t.313

::tz :;ii .27-f

5.613 0

Base .iqe at tat101 '4.133 8.466

Base ine at tatior 3.333 7.666

3ase Lne al tat101 3;ow c.333

Base B8.F.e lne at ine a tation tatio 7.466 0.800 1.799 5.133

Base Base ine 8' ine at itatio! tatior 16.667 0.000 51.000 4.333

0 .693 .953

__-_- 1.260 ;.;F

. 2 -e-m- 1.813 1.915 l.%O 1.959 1.935 1.883 1.800 1.667 smmm-

0 .640 .859

_--_- 1.155 1.360 1.58 m-e-- 1.688 1.785 1.828 1.827 1.800

:g 1:491

3ase he al ;atiox 5.667 LlOc4

Jase tne ai tat101

i:g

m--e- 1.483 ----- l.l.80 -m-s-

2; 4.432 0

Base\ Water line at statirJn L-l-- me1 2a&

14.133

II 0 0

.a33 .973

.167 1.210

.250 1.415 1532 ;.z;

,667 1:013 Loco 1.937 1.333 1.990 1.667 2.cco 2.000 1.990 2.333 1.957 2.667 1.903 3.0cO 1.827 ;.g; ;.;g;

4:cm 1:419 4.333 1.191 4.667 .873 4.833 .6'+1 4.917

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2.036

Base ater line ai lnel$$$

58.466

I.010 0 .167 .538 .a7 ----- 250 .663

.500 -----

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.WO 1.452

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.667 1.353

.@X.l.223

.333 1.015 '.5a, .wl .5m .767

'*667 2: ::g; 260 h92 ?857

atior .467 .800

-m--e 1.249

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1.124 1.237 .--se 1.387 L.471 L.502 k.4$

L:36o L.213

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0 .4oo .593 .815

1.067

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1.215 1.285 1.310 1.289 1.223 1.093

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4.0s: 0

0 .447 .607 .725 .827 .%I

l.loa L.150 LA3 t.363

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1:137 .879 .6-n .523 .250

3.7@J 0

0 2; .793 .m7

1.028 1.139 L.213 L.270 t.343 L.369 ~.363 1.323 1.235 1.030

.807

.573

.352 3.637 0

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.%3 1.065 1.147

::g 1.325 1.322 1.277 1.190 1.034

.674 A41

3.583 0

i.033 1.291 1.452 1.579 1.763 1.873 2.003 2.060 2.073 2.067 2.037

i:g g*g

. 1.510 1.313 1.033

$2 .627 .477 .227

5.190 0

3 .933

1.233 1.420 1.547 1.743 -A67 LO10 !.c83 ul.3 ,A03 2.030 2.037 L.963 ~,.870 1.743 L.583 1.363 1.073 .873 :TB: .373

5.140 0

0 .280 .497

:g; m--m- _---- 1.045 1.143 1.292 m---m

3.231 biius about water line 9.411

.709 aim

0 .343 .639 .9u

1.047

EC :893 .767 .f3r7 .565 .410

3.839 0

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1.383 1.416 1.4l.l 1.359 1.255 1.079

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hkster line IS distance above water line 6.6$7.

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TABLE III

PHYSICAL CHARACTERISTICS OF THE MODEL

Wing: Area(included),sqin. .............. Span,in. ..................... Aspect ,ratio ..................... Mean aerodynamic chord, in. . ., .......... Sweepback of leading edge, deg ........... Trailing edge sweep, deg .............. Incidence,deg . . ................. Airfoil section ................ ., .

. . . l . .

. . .

. . .

. . .

. . .

. . .

. . 987.26

. . 45.49

. . 2.096

. .

. . 2mg

i . -10

0064.08-g .NACA

Vertical tail: :Area,sqin .......................... 102.4

Aspect ratio ......................... 2.64 Sweepback of leading edge, deg ..... .,., ..... .... 52 Airfoil section .................... .. NACA 0005-64 Taper ratio ..... ., .......... ......... 0.324

Pod wing: Area (included), sq in. ................... 89.60 Span, in. .. ., ........................ 13.70

'Aspect ratio ......................... 2.10 Airfoil section .................... NACA 0004.5-64

Pod canard: Area(included).sqin.

.....................................................................

29.44 Span.in. 7.,86 Aspect ratio

.,.21 o

Airfoil section ........ .'. ......... WACA 0004.5-64

Pod ventral fin: Area (included), sq in. ..... ., ............ ., 19.20 Span, in. ............................ 4.10 Aspect ratio ......................... 1.75 Taper ratio

........................................ Airfoil section iAiA.oio4 ,".zz

Leading-edge sweep l -60 ....... ., ..............

L- -- 45.48 _/

---_-- ‘f:

77 -43 -.7--..Lp- 112-C

4.32-l t- . It 0

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(a) Sketch of complete model.

Figure la- General arrangement of model.. All dimensions are in inches.

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NACA RM SL56G23

M INUS -

-- ‘i

LOWER ORD.-

UPPER ORD.

- ._____

W ING.T. E.'

CHORD

*PLUS

,-CHORD PLANE

‘f- it tt B I--: , DIM.

f: AIRPLANE AND SYM.

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TANGENCY LINE1 W

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. - W ING STA. 0

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E

STA. 5.333 I #

SECTION E-E '2O13f45"

(b) W ing gemetry. (Also, see table I.)

Figure la- Continued.,

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NACA RM SL56@3

E SPAN STA., 5.333 1

t B

20+ I WING CONTOUR CHORD PLANE

OUTB'D CHORD PLANE -

SECTION G G

it SPAN STA. 5.333

d--J--- OUTB'D

LOOKiNG FW'D

SECTION H ii I

G I

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OUTB'D LOOKING FW 'D

E SPAN STA. 5.333

WING CONTOUR

. MAIN LANDING

I

GEAR FAIRING

TYPICAL SECTION IF DRAG STRUT FAIRING

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(c) Detail of drag strut fairing and niain landing-gear fairing..

Figure la- Continued.

t ;

WING STA. 2.792 I

9 -729

NAC. STA. 0

I WING STA. .2.747 WING STA. 16.859

CHORD LINE

FOUR BASE PRESSURE LOCATED 90° APART.

NAC. STA. 19.333

(a) Inboard engine nacelle and strut.

THRUST

PIC?cuPS

Figure la- Continued.

SPIKE SPAN STA.

17.333

WING STA.- 21.28% 30.022

CHORD PLANE

UST

NAC. STA. 0 19.334

(e) Outboard engine nacelle and strut.

.Figure l.- Continued.

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ii,

i,’ 8’. : .; / f 1. ,;

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NACA RM SL56G23

tt B”

I rTkl-mT.T.* rm i..Js”uuu- “- DMETRY

nmr tt A tt tt gt tt,,” 3.ui. , t - I

--- I nnn I I I YY7 I yy ’

i.007 I i.007 i -;BOo . A-w t .'I n,c I

. L,L,l, s.v*1

.170 1.019 1.019

.184 -

. 197 iii?-. ,

12.000 . 1.u.00 l.!.&.ooo 1.449

;.8;2 l Y Y

16,000 1.399 . __I^ 1 l . 621 . . -1 3Y 1a*:06: ?$7 A

18. 19.333 1:3W

1 1

(f) Nacelle geometry.

Figure l.- Continued.

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NACA RM SL56G23

F.. STA.

NAC. STA. 3 .h-00

0 .ooo

(g) Engine nacel le spike.

F'igure l.- Continued.

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31 ~ i. ;ii j. 3 i t+Y i f 1 1 ip

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NACA RM ~~56~23

\, STAR

WL 6.667

STA. STA. STA. 17.826 23.333 76 -467

l STRAIGHT LINE BETWEEN ORDINATES BEARING THIS SYMBOL r

STA. 0 STA. 1.00 STA. 6.000 STA. 10.000 STA. 12.159

WL1 BL WL1 ' L Ll BL Wl.l BL 'WL1 BL

1.04a 9 _.A~7 v,, ,... F.

lm 1~ distance above W I.6 -667 7'

(h) Fuselage geometry.

Fi@re l.- Continued.

‘Ci (6’ p,

3. 1

1.

i

I ’

I-.

NACA RM Sl3&23

E

FUSELAGE

BL "

.I

'

1WL is distance above W L 6.667

(i) Fuselage-wing fillet geometry.

Figure l.- Continued.

5 - ~-,.- .- -7,, .’ d =G

I

STA 8lij25 ii? TIP CHORD I

00

(j) Vertical fin.

Figure l.-.Continued.

NACA RM SL56G23

FAIR

-FAlRING

.TICAL FIN

1 ACTUATOR FAIRING SECrIONS I ---- I----- 1

s;rA.' Span 3tatlon 7.918 Dist. Dim A dius Fin&d

I I I

(k) Actuator fairing.

Figure l.- Continued.

y4

I 3 ;@.

i” :g. &i[’ 3, 3 ;ui :r 88 ii ‘i &.i, Y ‘a I ” ; j. ‘. 1 cjt : 4 ) .%

p1 1:

;q! ,d t 1% ‘# t .I!

1 */ . . r t,

i, I. 11 ,’ *

NACA RM SL56G23

RADOME GEOMETRY

STA.. 73.633 VE!RT. TAIL \ I

- RADOME

VERT. TAIL

RADOME STRUT

(2) Radome geometry.

Figure,l.-.Continued.

-- I

- . / . I-- ,L i

O R D . b

T-7- i

I-DISK-' / - -

(m) P o d a n d p o d strut g e o m e try.

F igu re l.- C o n tin u e d .

pa& > . . . /. .: --_- .- _ . ; W V ,#; ‘*$-a~+:

. ._.-- 7&

;-I-v-lrTTT*M .y~~~y~-~~~ f&&&~s&mA&? ;i & _ ~*~.&rdc

/

W ING STA. 16.859

A A. CHORD PLANE I;

D

. *-

.Ll. P

~

6 -ye&

lb,:' 1cI

(n) Strut geometry.

Figure l.- Continued.

$ ,C’,

t .*” L I’&

I,$

&f ,).c

1 f;,:: , & i’lr

:

I /I ,‘, :

i/

NACA RM SL%G23

POD STA. 2.120

POD 527 C

- .692

ORD. j---DIST.,-+ c I- - - b

I

TYPICAL SECTION N~A ooo4.p64

CANARD AT ZERO INCIDENCE AND DIHEDFML ROOT CHORD 2.200 INCHES BELOW PARTING PLANE NACA 0004.5-b4 AIRFOIL SECTION

I SECTION C-C 1 SECTION D-D CHORD DIST. \ ORD. 1 DIST. 1 ORD.

I I I I I I

- ._ - 15 1.124

'20 ---I 30 :%i

2:558 .16!~ 1 lb& I I080

40. .165 I Lh72 I 50 I I

, _~. 60 I 1.1

(0) Pod canard geometry.

Figure l.- Continued.

:

t E

POD STA.

POD STA.

TYPICAL SECTION- NACA ooo4.5-6L,

.T 4.100

(p) Pod ventral fin.

Fi&e l.- Continued.

I! I- ‘0 : 9, b’ ;

: > i’ ‘, 1 1 ‘i’r . j’

/ii

I’ ?, :$ f b 4;

:I

>2, i’ : 1’ a

! p

l.1’ : c :I. , *.. ‘, ‘. i :

Y 8 i ,,. .” I, .’ .

,d/

‘,, :_.

NACA RM SL76G23

POD STA . . ; 1 -F

16.000 A

1 - - T

TYPICAL SECTION NACA oool++6i+

I Pod Wing Geometry

Sect. A-A Sect. B-B Chord Mst. Ord. Dist. Ord.

I

(q) Pod wing geometry.

Figure l.- Concluded.

- ._..__~ -4~’ ;, -~’

-. ,. . ..‘Lf# ~_.. : ,-- i .:,. v. . ..hk&

I

--

(*

F?gure 2.~ photograph of model-

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; .; i “$a$ ..I, .S’,

I i ~ a ’ I. ^ )

: , ,. : - k

NACA RM ~~56~23

/I . ,

_’ ‘,*

-.., . . 1 ‘, :

‘-N, .‘*..,.L t, . *-

.’ ._ j,: .:\, i,“b

~-~=:T .I ;, “*:..$

I ;. ’ I$“&= _ _ --,-<*“., :,

t ,. .“i f __i ,‘;

: ;:: -I,, e-

i *&.

‘1 $I&

.,U’

L-92185.1 Figure 3.- Model,and booster in launch position.

(il .‘~- .p2&

‘. (_ ’ ,. ‘. . . . :

.6 ,8 1.0 1.2 1.4 1.6 1.8 2.0 2.2

M .

Figure 4.- Variation of Reynolds number with Mach number.

l$., \d .q

'DT

-. .

b.. I. -, .,. _ ;

.’ _.. : : ,.~.- -, ., ;*.$

.:

I ._.s ;<l;.,&~ T”.<

R I *

iii

Figure 5*- Total drag.

_ I

. . . -. :, :, c, ,:.

:’ . F,., :*, ..~L~~~ .i ii @.u 2% ,:L~~.Lr~i..&*c..;J;,~~~~~~~ 2

z *

Figure 6.- Base drag.

- =.- I. ,.$-a;

,- :

‘. : ’ 1

., I -, _ . . .

. . ,,,y: ..: .-,.....*‘.. _ a4 ” ~;~;&J+.~ 6;.;;&&!: 4 .~.i~~~~*~~~~~-~~~~i

Figure 7.- Internal drag.

, . - I ;_ i . . , -_ , ;-

.- Y . ) ; . I . ~ . ~” . .

: 1.. ,. :’ _ y., :. . -. l~.. ._<.. ;-. .-:‘,..y..~ .;.. ,/, .‘i .<*$,;“: :‘. . “.& . *. ‘: -’ .- ~ .,. : .,& & + ~ .~ ? ! p c ~ ~ ~ ~ ~ ~ ~ ‘-‘~ ~ ;l: ~ & & @ i $ & @ & ~ ~ ~ ~ ~

I

F

F igure 8 .- C o m p a r i s o n o f ex terna l d r a g d a ta a t m o d e l f l ight

a /& ‘7-T p 2 7 4 L m + + J a + < 4 froki p/v’ /O X 1 0 ’ +0 -6Ss /Oc

b “‘~ ‘+ ~ s ~ F//;bt fesf a A h ” ”

C n .0

r .I

: 1 . : . 3 I ., : &&& & .& ,. :1: yi ~~~~~~~~~~~~~~~~~~~~~ $ :. -,.:T--,, : . ._, I .:.

-.

*, c 0 %rrim

Figure 9.0 Trim normal force. Center of gravity at station 37.4.

. r

A!! ,’

Nondimensional cross-section area, A/S2 0 l g l Q b

‘F,. ~, '1 .P ;, ,I 7 ,,,J ', * ,: # . . .?* : .r 1 ,g;.j, .d$ f?" ..$$ ,<;i f/ f?' \g. 'jk' 2%. "$ 4 . 2 , $k xc e : 7 ' c&

.,I,,

,,

.I ,.

,:.;. ? “7

‘“2 -5 I /..

A’

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ig : 2;:~ g

f& 3

,$j

RM SL56G23 4 INDEX

A3 -;$\.

Subject

Airplanes - Specific Types Airplanes - Performance

Number

1.7.1.2 1.7.1.3

ABSTRACT

An investigation has been made by the Langley Pilotless Aircraft' Research Division utilizing a l/15-scale rocket-propelled model of the Convair ~-58 supersonic bomber. The drag at model trim lift was obtained at Mach numbers between 0.85 and 2.0 at corresponding Reynolds number per foot of 3.5 x lo6 and 13.7 x 106, respectively. The results of the pres- en-t investigation are compared with unpublished data obtained from several facilities, WADC lo-foot tunnel, Ames 6- by 6-foot supersonic tunnel and the Langley 16-foot transonic tunnel, A comparison of the drag at tran- sonic speeds and at approximately the same Reynolds numbers showed excel- lent agreement. A drag coefficient of 0.028 at a Mach number of 2.0 was obtained at zero-lift conditions.

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