Effects of landing gear, speed brake and protuberances on the longitudinal aerodynamic characteristics of an NASA supercritical-wing research airplane model

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OF

N83 11065

EFFECTS OF LANDING GEAR, -

(NASA-TM-X- 72684)

SPEED BEAKE AND PROTUBERANCES ON THE

LONGITUDINAL AEHODYNABIC CHARACTERISTICS OF

URClas AN NASA SUPERCRITICAL-WING RESEARCH AIRPLANE

A04/MF A01 CSC:L D.A 1 3-1; 0.2 3-2201

MODEL (NASA) 55 p HC

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NASA

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NASA Scientific and Technical Information Facility

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(NASA TECHNICAL

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NASA TM X-

MEMORANDUM

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y^ EFFECTS OF LANDING GEAR, SPEED BRAKE AND PROTUBERANCES ON THE LONGITUDINAL AERODYNAMIC " c CHARACTERISTICS OF AN NASA SUPERCRITICAL -Om"' ^.n WING RESEARCH AIRPLANE MODEL By Dennis W. Bartlett and Giul.i.ana Sangiorgio OE3 QUAL 1 t 5 Y,

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OF J A^ ^ Y 1 • -/7 'f • ^ • ^f NATIONAL AERONAUTICS hND SPACE ADMINISTRATION LANGLEY RESEARCH CENTER,. HAMPTON, VIRGINIA 73665 .r. 4, I Ifet>•sn Nu e 7 tio,mrrc , at Accession No 3 It. , . lm r.1 f, a o,•q TM X-72684 fit-.--_..._ n True and , S,rhatlo effects of Landing Gear, Speea Brake and 5 fiep'.1 July 1975 Protuberances on the Longitudinal Aerodynamic Charac t;,;lo teristics of an NASA Supercritical-Wing Research ar

Airplan e Model MO_

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t ^^ Dennis W. Bartlett and Giuliana*Sangiorgio _ 10 t w o u. t r,„ :1 Purlonnnnt Urganudr,n, 't 1, r• dml«nddn•t.

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MC1110-dndLe1 Technlcdl National Aerunautics Administration 14 SpUrlau4nq . r — u.y C:nk~ Washington, OL 20546 .^ .,. —..._...._,.._ .

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Special technical information release, not planned for formal NASA publication.

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Ain investigation has been conducted in the Langley Research Center 8-foot transonic pressure tunnel to determine the effects of thelI riding gear; 'speed brake and the major airplane protuberances on the longitudinal aerodynamic characteristics of an 0.087-scale model of the TF-8A super- F ' critical-wing research airplane. For the effects of the landing gear and z speed brake, tests were conducted at Mach numbers of 0.'25 and 0.35 with a J flap deflection of 20° and a horizontal-tail angle of -10". These conditions would simulate those required for take-off and landing. The effects of the protuberances were determined with the model configured for cruise (i.e., horizontal-tail angle of -2.5° and no other control deflection), and these tests were conducted at Mach numbers from 0.50 to ?.00. The angle-of-attack range for all tests varied from about -5° to 12°.

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PROTUBERANCES ON THE LONGITUDINAL AERODYNAMIC CHARACTERISTIC;', OF AN NASA SUPERCRITICAL- WING RESEARCH AIRPLANE MODEL By Dennis W. Bartlett and Giuliana Sangiorgio Langley Research Center SUMMARY An investigation has been conducted in the Langley Research Center 8-foot transonic pressure tunnel to determine the effocts of the landing gear, speed brake and the major airplane protuberances on the longitudinal aerodynamic characteristics of an 0.087-scale model of the TF-•8A supercritical-wing re- search'.airplane. For the effects of the landing gear and speed brake, tests were conducted at Mach numbers of 0.25 and 0.35 with a flap deflection of 200 and a horizontal-tail angle of -10 0 . These conditions would simulate those required • for.take-off and landing. The effects'of.the protubcranc!es'we ' re de- termined with the model configured for cruise (i.e. horizontal-tail angle of - 2,5 and no other control deflection), and these tests were conducted at Mach numbers from 0.50 to 1.00. The angle-of-attack range for all tests varied from about -5° to 120..

The extension of the landing gear resulted in a slight incrense'in lift and a small negative shift in pitching moment: (less than a comparable change in hori-.ontal-tail angle of 0.5 0 throughout the angle-of-attack range of the in- ) vestigation. The deployment of the speed brake (deflected 15 0 ), however, showed no appreciable effects on either the lift or pitching-moment characteristics.

a a As would be expected, the landing gear and speed brake did cause a,significant I I ri ^"—?""'^^- •^.+iw+n.s-+^—'^ ..

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.. . , _ ..«:+in..!*.»^,w.^ increase in drag, however, this increase was substantially larger near the minimum crag point (a ;z! 1°) than at the take-off and landing angle of attack of 8.5°.

The effect of the protuberances on the lift and pitching-moment charac- teristics is negligible, ,iowever, there is a small increase in drag throughout .

most of the lift-coefficient range at all Mach numbers. At the cruise lift coefficient of 0,4, the drag increment due to protuberances varies from about 0.0003 in drag coefficient at a Mach number of 0.50 to approximately 0.0008 at 0.95 Mach number. However, at the wing design Mach number of 0.99, the drag increment near 0.4 lift coefficient is about 0.0002 in drag coefficient.

INTRODUCTION In support of the flight-test program and simulator studies for the TF-8A supercritical-wing research airplane (ref. 1) and to establish the necessary data base for a correlation of wind-tunnel and flight data, extensive wind- tunnel tests have bAc--n conducted involving this airplane. In addition to con- figuration development-type programs (see refs. 2, 3, and 4 for example.), in- vestigations were performed to determine the basic longitudinal and lateral static stability characteristics (ref. 5), the dynamic stability characteristics 8). The (ref. 6) and wing and fuselage pressure distributionE (refs. 7 and purpose of this paper is to document the results of wind-tunnel tests that were conducted to determine the effects of the landing gear and speed brake on the longitudinal aerodynamic characteristics of the TF-8A supercritical-wing re- search airplane at Mach numbers near those for tale-off and landing (M = 0.25 and In addition, the effects of the ;ta,jor airplane protuberances (i.e.

0.35).

j antennae, nose probe, etc.) on the longitudinal aerodynamic characteristics are presented at Mach numbers from 0.50 to 1.00. Tests were conducted over an ORIGINAL PAGE IS OF POOR, QUALITY v .

POOR QuAU G Y,

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`s angle-of-attack range that varied from about -5 6 to about 12 0 and at Reynolds numbers which varied from approximately 10.2 x 10 6 per m (3.1 x 10 6 per ft) at 0.25 Mach number to a maximwn of about 20.0 x 10 6 per m (6.1 x 106 per ft) at 0.40 Mach number. Neat Mach 1.0, the test Reynolds number was about 16.0 x 106 per m (4.9 x 10 per ft).

SYMBOLS The longitudinal aerodynamic characteristics presented herein are referred ^.o the stability axis system. Force and moment data have been reduced to , con- ventional coefficient form based on the geometry of the reference wing planform, which is the planform produced by extending the straight leading and trailing edges of the outboard sections of the wing to the fuselage center line. (See fig. 1(a).) Moments are referenced to the quarter-chord point (fuselage station E cm (3915,5 in.) ) of the mean geometric chord of the reference wing p"anel.

99•45

All dimensional values are given in both SI and U.S. Customary Units; however, measurem:ntz and calculations were made in U.S. Customary Units.

Coefficient's and symbols used herein are • defined'as follows: b wing span, 114.30 centimeters (45.00 inches) Dz Sg CD drag coefficient, q Lift lift coefficient, CL q Pitchint^ moment Cm pitching-mcinent coefficient, 9^c c mean geometric chord of reference wing panel, 18.09 centimeters (7.121 inches) local streamwise chord of wing c free-'stream Mach number M free-stream dynamic pressure q

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0.193 meter (2.075 feet`)

y I 1 i a angle of attack, referred to a model water line, degrees horizontal tail deflection angle, referred to model water: line 6h (positive when trailing edge is clown), degrees

s

flap deflection angle (positive when trailing edge is down), degrees TEST FACILITY

The investigation was conducted in the Langley Research Center 8-foot

transonic pressure tunnel (ref.

9). This facility is a continuous-flow, single-

return, rectangular slotted-throat tunnel having controls that allow for the independent variation of Mach number, density, stagnation temperature and dew- point. The test section is square in cross section with the upper and lower i

walls ax ally -slotted (each wall having an open ratio of approxi=mately 0.06) to

permit changing the test-section Mach number continuously through the transonic speed range. The stagnation pressure in the tunnel can be varied

from a minimum

value of about 0.25 atmosphere at all test Mach numbers to m'aximirm valuL-•of a approximately 1..5 atmospheres at transonic Mach numbers and approximately 2.0 atmospheres at Mach numbers of 0.40 or less.

MODEL DESCRIPTION Geometric characteristics of the

0.087 -scale research airplane model are

presented in figure 1, and photographs of the model are presented as figure 2.

The basic fuselage and tails are scaled versions of those utilized on the test- bed airplane (TF-8A). The model ws equipped with flow-through ducts which discharge at the base of the fuselage on either side of the flat-sided model support sting. Internal drag coefficients and mass-flow ratios are contained in ref. 5.

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The wind; used during the investigation to determine the effects of the landing gear and speed brake was constructed of aluminum. A flap deflection of 20 0 was employed for these tests with a horizontal-tail angle of -100.

r" These flap and tail angles are those that were estimated to be required for take-off and landing. Flap and aileron control-effectiveness data obtained with this aluminum "control" win,a are contained in references 10 and 11. To obtain wind-tunnel performance and pressure data for the research airplane, a separate steel wing was normally employed (see ref. 2 for example), and it was the steel wing that was used during the present tests to determine the effects of the major airplane protuberances. Both win,7a are geometrically the same, and coordinates are presented in reference 2.

Ir The supercritical wing was mounted on the fuselage at a root-chord in-

i'

^j cidence angle of 1.5° and has approximately 5 0 of twist (washout) from root to is tip in the'unloaded condition.. The reference wing planform, which excludes the leading-edge glove and trailing-edge extension, has a taper ratio of 0.36, j' an aspect ratio of 6.8, and h2.2'' O of sweepback at the quart er-chord,line.- The I^ area of the reference wing planform including the fuselage interce pt is 0.193 m`" ix t (2.075•ft 2 ), and the mean geometric chord of the reference wing panel is 18.09 cm i^ (7.121 in.).

Detail.; of the model landing gear and speed brake are presented in figures 1(,b), 1(c) and l(d), and the landin„ gear and speed brake are shown on the model in the photographs of figure 2. The basic aircraft speed brake, deflected approximately 15°, was used during landing to aid in stopping the airplane which was not provided with a. drogue chute. Details of' the major airNl•tn,.

protuberances ' are presented in figure 1(e), and these protuberance.; are.also shown in the photographs of figure 2.

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Y x li ti..a b .,^^ '. r:3 OF PQ11^^a9 d Y The underwing, leading;-edL,c vortex generators (fig. l(f)) were emplaycd- on both model wings (0-0-percent-wing-semispan station) for all tests of the .

i i present investigation, however, the aileron hinge fairings (fig. 1(g) and 1(h)) i t were included only on the steel wing. The underwing, leading;-edge vortex gen- erators are discussed in references 3 and 12, and limited results for the effects of the aileron hinge fairings are also presented in reference MEASUREMENT y AND TEST CONDITIONS Measurements,of overall forces and moments on the model were obtained from a six-component, electrical strain-gage balance housed within the fuselu^e cavity. Differential pressure transducers referenced to free-stream static pressure were used to measure the pressure in the fuselage balance chamber and at the model base.

The effects of the landing gear and speed brake were measured at Mach .

numbers of 0.25 and 0.35 for a flap setting of 20° and a horizontal-tail angle of -10°. For determination of the protuberance effects, the model had no 2.5 0 (estimated to_ control deflebti.on other than a horizontal-tail.angle of - be that required for trim at the design-cruise condition), and measurements were obtained at Mach numbers from 0.50 to 1.00. The angle-of-attack range for all tests varied from about -5° to approximately 12° for a sideslip angle The tunnel. test conditions at the !•loch numbers of the present investi- of 00 .

gation are presented in table I.

Boundary-Layer Transition The boundary-layer trip arrangements used for the wing are shown in No. 120 Carborundum grains were located on the horizontal and figure 3.

ver.:' , ..1 tails at 5 percent of the local streamwise chords and were also applied 2.54 cm (1.j0 in.) aft of the model nose and 1.27 cm (0.50 in.)

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OF p00'3 QI9AMY

rearward of the inlet lip on both the inner and outer surface,. All boundary- layer trips were applied to the model in bands that were 0.1 27 cm (0.05 in.)

wide and were located by neasur • ements, taken in the streamwise direction.

Corrections Drag coefficients contained herein have been adjusted to correspond to a condition of free-stream static pr • essuro acting; in the balance chamber and at the model base (excluding the duct exit area). No adjustments have been made i to the drag, however, for internal duct drag. (See ref. 5.)

Corrections have been made to the measured angles of attack to account for deflecti.on' of the model balance and sting cupport system under aerodynamic load and for tunnel airflow angularity.

PRESEIITATIOPI OF RESULT The repults of this investi gation are presented in the following figures: i Figure r ' Effects of landing; gear and speed brake on longituainal aerodynamic characteristics. 6 0 ; ; f = 20° = -10 b i r Effect of protuberances on longitudinal aerodynamic chaacteristics.

I

. .

2.50; . . . . . . . . . . . . . . . . 5

= 6f = 0 0 . . . . .

6h - fVariation with Mach number of the drag increment due to protuberances DISCUSSION OF RESULTS, Effects of Landing, Gear and Speed Brake The effects of the landing, gear and the speed brake on the longitudinal aerodynamic characteristics are presented in figure 4 at Much numbers 0..25 and r0.35 (take-off and landing Mach number range). From this .figure it can be seen that the extension of the landing gear causes a small increaue in lift t;hr•ough- out the angle-of-attack range. This is probably due to the fact that the r..

i);iI{:IXAL NWE IS _^' POOR QUALtl`Y I relatively large plain fear doors (fig. 1(c)), which form an angle of about 200 with the horizontal in the deployed state, are producing lift. In addition, the deployment of the landing gear also results in a small negative shift in pitching, moment (ACm 0.015) through put• the majority of the angle-uf-attack ranee, however, this only amounts to a comparable change in horizontal-tail .

angle of less than 0.5 0 . The speed brake (deflected -- 0 ) fi,;,,s only negligible effects on the lift and the pitching-moment characteristics. (See fig. h.)

Near minimum drag (a ti 1°), the landing gear causes an increase in drag of approximately 50 percent over the basic configuration, while the speed brake results in about a 23' pe.•cent drag increase. However, near the angle of attach for take-off and landing (a = 8.59, the drag increase resulting from the landing gear and speed brake is.about 29 percent and 11 percent respectively.

As would be expected, there is little variation in the effects noted above between the two Mach numbers at which data are presented in figure h (M = 0.25 and 0.35).

Effect: of Protuberances The term "protuberances", as used in this report, includes the airspeed probe, the camera fairing plate, the PC14 antenna, the anticollision light and, the drain valve (fig. 1(e)).

The cffect of the protuberances on the longitudinal aerodynamic character- istics is presented in figure at Mach numbers from 0.5 to 1.0, and the drag 6 at lift coefficients of increment due to protuberances is presented in figure 0.1 and 0.4.

The protuberances have no appreciable effect on the lift and pitching moment characterirtics, however, as would be expected., there is a small increase in drag throughout most of the lift-coefficient range at all Mach numbers i t r i

r^

r j presented in figure 5. At the cruise lift coefficient of 0.4, the protuber- ance drag increment varies from about 0.0003 in drag coefficient at a Mach number 0.50 to about 0.0006 at 0.95 Mach number. However, at the wing design . Mach -umber of 0.99, the drag $.nererent due to protuberances near 0.4 lift r coefficient is about 0.0002 in drab; coefficient (See fig. 6.) Similar effects are also noted in figure 6 at a 'lift coefficient of 0:l which is near s s ! minimum drag.

` I The drag increments due to surface defects (slots, gaps, scratches, etc.)

for the basic rF-8A airplane ere documented in reference 13. These drag in- cremenus along with those reported herein must, of course, be consideredrwhen attempting to extrapolate wind-tunnel derived drag data to full-scale condi- s tions. (See for example, paper 5 of ref. 1.)

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u ^' REFERENCE; 1. Anon: Supercritical Wing Technology - A Progress R_port on Flight evaluations. NAGk uP-301, 1972.

2.

Bartlett, Dennis W.; and Harris, Charles D.: Aerodynamic C'"aracterL'stics of an NAOA Supercritical-Wint; Research Airplane Model With and Without Fuselage Area-Rule Additions at Mach 0.25 to 1.00. NASA TI .f X-2633, 1972.

Bartlett, Dennis W.; Harris, Charles D.; and Kelly, Thomas C.: Wind-Tunnel 3.

Developnent of Underwing Leading-Edge Vortex Generators on an NASA Super- critical-Wing Research Airplane Configuration. NASA T14 X-2808, 1973.

4. Bartlett, Dennis W.; and Harris, Charles D.: Effects of Wing Trailing-Edge Truncation on the Aerodynamic Characteristics of an NASA Supercritical- Wing Research Airplane Model. htASA TM X-3024, 1974.

5. Bartlett, .Dennis W.; and Re, Richard J.: Investigation cif

Basic Aerodynamic Characteristics of a Fupercritical-Wing Research Air- plane Configuration. NASA TM X-2470, 1972.

I itch, Yaw, and 6. Boyder., Richmond P.: Dynamic-Stability Characteristic:• in P NASA TM Y.-2900, Roll of a Supercritical Wing Research Airplane Model.

i L 1974.

!

7. Harris, Charles D.: Wind-Tunnel Measurements of Aerodynamic Load Distribu- t r tion on an NASA Supercritical-Wing Research Airplane Configuration.. NASA TM X- 2469, 1971.

Harris, Charles D.; and Bartlett, Dennis W.: Tabulated Pressure Measure- 8.

r mentc on an NASA Supercritical-Wing Research Airplane Model With and Without Fuselage Area-Rule Additions at Mach 0.25 to 1.00. NASA ^' TM X-2634, 1972.

i 9. Sebwvfer, William T., Jr.: Characteristics of Major Active Wind 2lmneIa t at Langley Research Center. NASA TM X-1130, 1965.

• 10. Bartlett, Dennis W.: Effects of Differential and Symmetrical Aileron De- fle,tion on th Longitudinal and Lateral Aerodynamic Characteristics of an NASA Supercritical-Wind; Research Airplane I 4,odel. NASA TM X-3231, s 1975.

d l]. Be, Richard J.: Stability and Control Characteristics,, Including Aileron Hinge Laments, of a Model of a Supercritical-Wing Research Airplane.

UASA T4•X-2929, 1974.

F 12. Harris, Charles D.; and Bartlett, Dennis W.: 111nd-Tunnel Investigation of i Effects of Underwing Leading-Edge Vortex Generators on a Supercritical- , !

Wing Research Airplane Configuration. NASA TM X- 21171, 1972., 13. Horton; "Uti:r A.; and Tetervin, Neel: Measured Surface Defects on Typical i Transonic Airplanes and Analysis of Their Drag Contributions, NASA TN D-1024, 1962.

i a ORIGINAL PAGE IS (h^ roolt QUALITY F w k I E I F i • Y I i M L it roes OF PGOR QJ4%I—"9`VV TABLE I.- TWIMEL TEST CONDITIONS Temperature Reynolds number Macb number K OF perm per ft

1.00 322 120 106 4.8 x 106

15.9 x 4.9

322 120 16.1

•99 4.9

.98 322 120 16.1

4.9 322 120 16.1 .97

322 120 16.11 5.0

.95

.90 322 5.6

120 18.4

.80 6.1

322 120 I 20.0

322 120 1 4 4 4.4

.50

322 120 7.2 2.2

.35.

.25 322 120 10.2 3.1

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( 23.8719,3961 - iusebge slolwn 130.13151.234) 1 l.tadwq edye and 4 ,10..q COW ti^ 1 [ of fne case an d panel ••— 3,531! 3681 $.18,0917,121) - — f . 114.30145,00^!

I , 9.00(,3.545) Model fuselage sfafm 2 8.711 11.305 ) 21.42 99.45139.155) 32 02112.6081 (8.435) - J .44(,174) 'Nolen 22.09(8.69.1) • , •r _ ,1 t 161.39!63 54) (a) General arrangement of 0.087-scale model.

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Figure I.— Model details. Dimensions are given in centimeters (inches).

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Fuselamp station 71.56(28,17) "r ^- — Model now 1 --- Bottom of fuselage 1.99L 78) i I PCM antenna i (Located on bottom center line of fuselage) + .8805) —^^ 171.07) Fuselage station 19.13(31.39) Fuselage station 81. 94(32.26) 55LKI — -2.21(.87) — r— t 661.26) Side view ront view Anticollision light (Located 15.24 cm 16in,) left of bottom center line of fuselage) Fuselage station 137.05(53. 96) Fuselage center line Left side of fuselage I \ 2,32(.9 1) 18lmax.l (. 01) '} Fuselage station _ Drain valve 3 129.98(51, 17) `^' 1.99L 78)- 1.991.18) I_ . j i Drain valve S. 1412.2b) ' -f -- ^ ' 18°

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(e) Simulated full-scale airplane protuberances. Concluded.

Figure I.- Continued.

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r

N

t N `n (3) M u

> i

C O • CO) LA- cu • p tf) N

d, o v

1 U ( ^ U s ^ 1 ^ L a Q1 '- C - ► - X a • 0 O ]_A E 1 ^r

r +ow

i ^ .

. ,...t i i 1. „ .,., ^ _. ! _Y , _ _... _. .. { r ^ _. _^..•-....`,...-,.i _ ^.,,,J,,,,.

4x . .:i'•;Ltiii7^-- ^L '"""^IIL'. '"'1^'IS^aCL:' 1'.fl?:^'1Lr1Y^ * ^'_f!lT."+.1i,^"`,Lf"s^ ^ L...'S'1^" ..^'Si3^:21,' ) r"^5 ^ .CX'S'^ t^^.^.^^' 7i'.^Jt[^r^^.^ ^^la^^ u Y OR V .

OF PO i I a ^L w O n O ^ C

d ^

^ O

O U

^ I O Cu C• ^ o Gi a. ^- u ' O J' Q7 p AGr IS ORIGINA^, It 4t^L11Y ^` Ol PUU • - i

i

x 20 ^ *. ^ ^. i ^ i ..1. I ._ J.... 1 J .. .^ ^^,,.a-,^^!... J. x.11..._.-•-..- .Y' .....

......5.._...^^^c.w.r^s`...^ • R ^ ' L', - _•___ ^..

r I

}

I

I

of

r Pion view J

i

f W,ng lower surface iElewion , I , ^! ^O Cross seclions s .i 't r: { ?^ (h) Cross section of aileron hinge fairings.

Figure I.- Concluded.

r i t .

r

of

jj W..

rr

ORIGINAL PAGE IS OF POOR •-'QUALITY „^:X^G k Y

n 1 KIM At OF i`O1U4yW

I I t .{ R w, y^M .t. X '” μ K.rtiF ' kt i ^^.. X a • l `, rr.'$±' ek yV'i 7 r «. e• fX c.t Y7 i I i '% w 7 A 1 e

4^

1-. [

z r !

X

711r. ^rl

^' ` bue., ` '` ` t ' r.c.'eaE XF ^, f ^ 1 ^ s ^ ^'! '^^tri 'r ^J ctr g-r -^"-•a^^!i •3 ` rX r IIASA L-11•L991 ^^ ^ Concluded.

Fh ute 2.

^.:Jsltl^at^ai^caZnrd•' 1 A T d 31t:

OF ^» ••

e No 100corborundum groins 28 50' 5) J 12 No.90 Side of fuselage r.. \ ^ `,gyp ♦5 ♦ --- yL odd 19051750) Wing lower-surface arra ngement y^`° No 120 ♦.

Used at all )doh nurnbeis' r 19.05(750) No 100 No.100 2286(900) ♦ - .05C NO .102858(11 25) 2032(8 00) r ' , f ' 1346(530) .508 (2 00) r

i 1

Fuselage Wing upper-surface rearward arrangement Wed of Mr0.95 to 1.00 ^ t

I lJ ( To wing tip

05C - r .05C --^' No.= .05C - 21 '59(850) i 1588(625) ., 1.

^ p 05C — No 10v 1016(4-00) _.

508(200) f Wing upper-surface forward arrangewnt i Used of M.025 to 0.90 Figure 3.- Wing boundary-layer trip arrangements. Dimensions are in cm (in.).

,

s

t

^ ., i -f '# j t j t ^.

^.

' r ^.

.` , ^'^ '.

^:

.1a

Speed broke land. gear i 0 Off Off f ,3 ( f 6 0 Off On On On .3 .2 l 1 _^.

^_ _;—

.z4

Cm .16 i .12 .08 — .0a —,04, ti b t3 10 12 14 16 I I S • a,deg

(a)

M = 0.25. Concluded uded

(

Figure 4.- Continued.

f3P YOUR QUA1=1 # i. v

i

.-.. .. _ .... _...Y . r---• - y_«ter—.

^ - ,.,^ .06 .04 A2 f 1 I ^ I i

L- O

LL

-12 -10 -6 -6 -4 -2 0 2 4 6 8 10 12 14 16 a,deq E (b) M = 0.35. Continued, • Figure 4.- Continued.

R

i

?.9

it t t 1^ .j I ^c - sg 1,2 0^ 1.1 ^ a n off I, I^ - I f s i • 11 + CL S .2 f r • , i s • S •6 -+ • S -2 —I O I 2 S + 10 II S 6 7 B 12 13 I+ 9 IS 16 (a)M=0.50.

Figure 5.- Effed of protuberances on longitudinal aerodynamic

characteristics, bh = - 2.5°; bf = 01.

<< ORIGINAL PAGE IS OF POOR QUALITY it k

j a

ff"71

.20

o

On off .19 .16 .16 .15 .13 .12 .11 co .07 . . . ... .........

.06 .05 , 04 '03 . . ........ .

.01 9 10 II 3 .4 5 6 .7 5 4 3 1 0 1 2 L 0 61 1 C-i 7 M., k , I - p I n , C ., 1 ;77 1s

OF POOR QU, ,IILIYV

= 0.50. Continued.

(a) M

-A

Figure 5. Uum "U

10 11 6 7 a 9

OPIC

OF POOR QUALITY

is .4 .5 .6 .7 .8 9 it itinued.

ORIGINAL PAGE W OF POOR QUALITY wed.

n T °^ era r,-^7 4 ^^: ^^"±^ OF FOP, QUALITY ....

^Peotuberance5 .........

On 1 ».}. ., ..t Orr, .., ...a . { ,16 .12 • i b c m -04 —08 —,z0 -.24 _ .. 1 .......

-20 -.6 -.5 -4 -3 • -2 -.I 0 I .2 .3 10 .4 5 .6 .7 .0 9 i I (b) M = 0.80. Con cl uded.

i Figure 5.- Continued.

i, y a t, tr.r "p' f .: ^.LA..

"`.`;'Y''3`r^ ,:r.!^^^lr^r'r"'j^r^xY^^^.^^^^..-L„^....1.^.,.«^f'^».1.,...,..,L^.^..',»1.,»^..L'. ^1 ► ^...y.1^^^.^:.rrJ•^t f Proroberonces „» t. 1 _ On 7 - 1,0 o Off --!

. t ,9 • i ce ^. a. ^.

1 t,• .^• I . ^ iii, I ..»..•I .8 I 1 ^ .7 .6 ,5 .4 .3 C L .2 a i 'I } -.2 i -.4 _ ..i...,..

', -,s —.

-,6 -6 -5 -4 -3 -2 -1 0 1 2 3 4 5 6 10 11 7 8 9 *,deg (C) M = 0.90.

PACE' 15 oRI mI Ai- Wit QUAL11Mi r^ ( or p 1 } Figure 5.- Continued, t;

i

1'.

I:

^^^^ I

or

i Protubsronces I i On

i

.I 6 .., _...

0 Off

. I

.1 .I f I .

.I 1

i

.10 .0s

c

0 08 .05

r

I •.0a 03 .. I ....

C r T Ll —1-- .01

i

-,6 -.5 -A -,3 -2 0 1 .2 3 5 I0 A 6 74 9 CL r s

(c) M

0.90. Continued.

Figure 5.- Continued. r

r j E F 1 ^ ^+ Pr0lubUf01KCS 1 on a - t ^ .

a .20 .16 e .12 .08 i 00 —.04 Cm i —08 f t + _ ^.i.. ._: . I t .. s —.16 i a^.

—.20 1 .

—.24 1 —.28 - .6 -.5 —.4 —.3 —.2 10 —.1 0 .1 .2 3 .4 5 .6 .7 ..8 9 CL ems ; (c) M = 0.90, Concluded.

OR.IGLNIAT PAGE IS

.S {

Figure 5. -

Continued.

! OF POUR QUALITY ' L r _ ,t

OF

. ... . .....

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

}..

» _ 1_ _ ..^»._...^ CO .05 ^..^. 04 -

: I

.03 ................

--- --- 02 ......... ...

.... .. . ....

0 1 A.... I.

IJI-

0 .....

.3 .4 .5 .6 .7 1.0 c L Ip (d) M 0.95.

Continued.

X

VA w

Figure O 5. • Continued.

= ......... .

rnnlr^f^i.LltJ.d1^_ °^'^ ! Ploluberonces ... !. ... I. .... ... .:.. .....

f U 011 i ^i...^ ..

i

f ' { .......... ......^. , . ..

.... _.:._

... .^

f.. :... '.

.12

.08

I. .....r.

.....^...... ......^ . ....... ...... _.^. ......,._...

.j..,... i..

-04 -08 : ..0 ...

i ,... : f .......... ...

... '

.. ......^...............:.. ' .:.. ;......

-.20 :. .

I . ^. .. .

t L..., +. .,.

-.24 r.

:..

...l..........

-,6 -.5 -.4 -.3 -.2 -.I 0 I 2 3 4 5 6 7 a 9 10 CL (d) M = 0.95. Concluded.

Figure 5.- Continued.

d

I

t —i Prolubcrunces i j On im ^. ... .. .... ....

:. q .. .... . '.. ....:. r .....^ :.,..

..,

j. } . }

.7 _... ! y. ^...... I ,. { ...

_.. ^...... _ ..:..

l . ...• .^ } .

i

I !

VI ^ 1 I ^ -5 -4 -3 -2 -1 0 1 2 3 4 5 G 7 8 3 10 it a, deg •

O..^t'^

OF POU'a .

(e) M = 0.97.

Figure 5. - Continued.

:1 ^{ r.....r ` j I !. L. ' Rl' .• C ' .Fl+ r. ...1:. . •I.

i !

OF POU

I

i 19 Protuberances On i 0 I !

R j »l. _i.... ... .. .... ....

.18 i.

a Off j .. .. t .. ... , .^.........

I 17^ .a ..... ..... .... .. _... .... .,.. ...» .^. ..,, .I .... a ....^ ••-^, ^ .16 S .._.. .. _. + ..

13^ .....^.,,. _.» .... ......... _.. ..... .._r._.. »j.. ».._ .,....l......_.

,12I ,. ....

10f j ,^ 06 J .04 .......

;.

Y .03 I .01

i

^• """ ^^ra1h• r t `' , • [i .....

!

t•,olubcrunces t . ...

. I ....

..

.^.

r . .. .i .....

U Ott ! A `h to 1 32 x t ^ ^ ^^ i I 1 i I s ( 5 ;.

20 i { .^.

..........

^..

161 I I.... i i ^... I ,.

t ... i ^\ ^..... l t 2 ^...... f......

.... .....

i I^; 09 t 1 ; I I ^ 1 , 1 I ' r !

I 1 i I t C, o i s (.....}, -of) i ;

.v

-24 f - •6 K -4 -3 -.2 -.I 0 1 2 3 .a 10 1.1 S 6 7 N 9 CL M = 0.97. Concluded.

(e)

Figure 5.- Continued.

`

i ^ Mr c irl-mt'- !

" On ^ Off ~ / ' | .

, /.

r a ^y « .3 .2 ^

C L '/

| ^ o ^ `/ i / ^ ^ -.o ' ' '» i ^ ^ .| ^ ~ ~ ' u ' s .o ~* + -z o / o s * u o r o o c.dee

IS'

ORIGINAL PAGE

^ Figure 5.- Continued.

QUAIIM

or , POOR

> .

| .

/ ..

..

.

.

' !

.

^ ^ ^( ^' ^ ]___] ^ ^ .

.

^ [ ' ' !

^ 'r-1,01'Utwronces .16 Off '06 c, —04 , r —08 —12 .. ........ ......... . .. . .. ....

.06 .05 .04 ltt U3 .04 I . , I ... .... . ... ..... ........f ... : :. ._:..

^. 1

lot -.1 -,3 -.2 0 1 .2 3 .4 .5 6 .7 a 9 10 CL (f) M 0.98. Con cl uded.

Figure 5.- Continued.

A :Z:z c^ d0MmrwrrrTVLTVI I tA L

OF

h- II Protubvrances Oil I--- ---i— —t Oil Lo • Pf ... . .....

...

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

... --4 ........ ..... ...

1 3 4 5 6 7 8 9 a deg (g) M = 0. 99.

gure 5. - Continued.

ii r^ —P- :ICIIICAMAI_ •r ..v J I$otuberuncec On 7 I .. _ _ ...

!.. _ .. . —. ^... ..... r

I

a off k

.I

.i .I I r i }!

i

i .11

i

CD i f .08 r .OS ,03 ^^ I AI -.4 -.3 -.2 -.I 0 I .2 3 4 .5 6 .7 a 10 ! I DL t ORIGINAL PAGr" I5 (g) M _ 0.99. Continued.,

Qu"rry

Pooa

OF

Figure 5. - Continued.

A 1

_.. . rF.-- ^^...

I

- | MFPG",a)^^^i_^u.

| ^.

, ' PrOluberorKes ^ On ' ^ .

.2u / ^ ^ / IN xo ./u ^ / ./e i ^ oo ' ow C'

IN

o -o* _oo' / — /c . | i | _/o | ' ^ -201 / -o^ .... ..... . . .. .... ... ... .......

-ou ..

` u '* '3 .3

. +s `o . -J o / x ^ o o .7 o to

y I CL ` ` Cn\ KA 0.99, COOC!Ud8d.

.

. '.

" ' Figure 5.- Continued.

' .

^ ' ~ , . .

^ ' ~ ' ^ ^O .. ---^^r .

^ .. ' ^ ^ ~ ' | ^ | ^ u , On ~ " IX ^ | ` ~ c / / ` ~^ | X ^ X / .

.` -. ._ .

CLu !

Al

' u x ' u -,I ^ ^ -« -.» .^^ ^.4 .|.

^ '^ / ' '5 ^ , ^ '16 .` .

K^ (h\ K8 ^^ ^ x^l^^` 1. OO --^ `^' ^' i^ ,_ .

Figure 5.- Continued.

^ ~ .

.

5!

.

.

..

OF

- ... - , " — - —:— " "" T Prolubtronces: on Ott .16 C -04 m • -08 -12 .07 .

.... ......... . .. ......... . . ..... ..... . . '06 ,05 .. ........ . 04 t.

of 1+ -. 5 -4 -.3 -.1 0 .3 A 5 6 7 0 9 -.2 .1 .2 CL • (h) M 1.00. Concluded.

Figure 5.- Concluded.

tt

t• ^e .002 (AC D^ C (D.1 -----,--

p - -- -^

!

r _ 4.

I

C

(ACID)p = D, protuberances cri CD,protuberanre

J b'02- l (OCD).

C) .4 6 T c^ ,`;

• M

Figure 6.- Variation with Mach number of the, to protUberanc,es.

ORIGINAL PAGE LS OF POUR QUU j ^1. • (`jrr r ti.•^''•^.. •^ ,.^..,. ...

j'''am ^..^r .'.. .....«.^........ ^.^ ...^._.. ... ^.. ,... _ ,-.....-^---...,.. .., ..... Yom. ...

.. ,.• ,• ..

..........._.. __ Jam .

.^..

i

y.^,.-. ^.

-• ..r,,,...._r ^ ;,,,,^,^,,,._.

....__..__ _.._... ._ ..

-- .. ^ - — ^-- - - j