Lift and Drag of a Swept-Wing Fighter Airplane at Transonic and Supersonic Speeds

- Restriction/Classification

2 7 0 0 0 3 copy 383

Cancelled c C I .

MEMORANDUM

LIFT AND DRAG O F A SWEPT-WING FIGHTER AIRSLANE AT TRANSONIC AND SUPERSONIC SPEEDS By Jack Nugent High-speed Flight Station Edwards, Calif.

C A S

C

Restriction/Classification Cancelled TU8 matertal contab lnformauon ths m e w of thn e s p l o x v ~ laws, T i t l e 18, U.S ~ f which in slly manner t o an unnuthorized p n o n Ls pmhiblted by l a w .

NATIONAL AERONAUTICS AND

SPACE ADMINISTRATION

WASHINGTON January 1959 Restriction/

CON

Classification Cancelled R NATIONAL AERONAUTICS AND SPACE ADMINISTRATION r mom mu^ 10-1-58~ LIFT AND DRAG OF A SWEPT-WING FIGHTER AIRPLANE AT TRANSONIC AND SUPEESONIC SPEEDS" By Jack Nugent SUMMARY A flight investigation was made of the lift and drag of a swept- wing fighter airplane in the basic configuration and in a slats-locked- closed configuration over a Mach number range from about 0.63 to about 1.44.

At a nominal lift coefficient of 0 . 1 negligible drag-coefficient difference existed between the two configurations over a comparable Mach .

number and altitude range. For the basic configuration at zero lift the supersonic drag level was about three times as great as the subsonic drag level, which was about 0.01, whereas the drag-due-to-lift factor increased about 137 percent over the test Mach number range. At comparable Mach numbers the high-altitude data produced a larger lift-curve slope and showed a more pronounced variation of lift-curve slope in the transonic region than did the low-altitude data. For the high-altitude data the lift-curve slope at a Mach number of 1.44 was approximately 62 percent of the value at a Mach number of 0.9.

INTRODUCTION In recent years the NASA High-speed Flight Station, Edwards, Calif., has conducted in-flight lift and drag measurements on current airplanes as part of the joint Air Force-Navy-NASA high-speed flight research pro- gram. This paper presents the results of flight tests of a swept-wing fighter airplane with speed capabilities varying from subsonic to w e l l within the supersonic region. Lift and drag measured for the basic config- uration (free-floating wing leading-edge slats) are presented. The Mach number range extended from about 0.63 to about 1.44 over the usable lift range of the airplane.

Data were obtained over an altitude range from about 20,000 to about 40,000 feet during push-down turn maneuvers and "Title, Unclassified.

Restriction/Classification Cancelled C OW I accelerated maneuvers; limited data were obtained from speed runs.

Data for a configuration in which the slats were locked closed for all flight conditions also were obtained concurrently with a maneuvering- characteristics program. The Mach number range of these tests extended from about 0.87 to about 1.13 over the usable lift range of the airplane.

SYMBOLS A aspect ratio, or cross-sectional area, sq ft inlet duct area at pressure-measuring station, sq ft Ad exit area of jet nozzle measured cold, sq ft ' e an measured normal acceleration, g units measured longitudinal acceleration, g units airplane aerodynamic drag coefficient .

CD dCD

-

drag-due-to-lift factor d C 2

c f turbojet nozzle coefficient

CL airplane aerodynamic lift coefficient -1 slope of lift curve, d e g ' l , radians C La C wing chord, ft

-

wing mean aerodynamic chord, ft C gross thrust, lb Fj

Fn net thrust, Fj - Fr, lb

Restriction/Classification CONF Cancelled . . . . . . . .........................

Fr ram drag, lb acceleration due to gravity, ft/sec 2 g pressure altitude, ft hP local slope of engine thrust-rotational speed curve, lb/rpm kl 6 ,

constant for flight at a given Mach number, k -, sq ft/rpm

k2 Pa 2 configuration length, ft maximum value of lift-drag ratio (L/D)mx M airplane Mach number inlet-duct Mach number at pressure-measuring station N 1 low-speed rotor, revolutions per minute ambient static pressure, lb/sq ft Pa inlet-duct static pressure at pressure-measuring station, pd lb/sq ft total pressure at compressor face, lb/sq ft P' t o t a l pressure near jet-nozzle exit, lb/sq ft p'e 9 free-stream dynamic pressure, 0.7M2pa, lb/sq ft S wing area, sq ft total temperature at compressor face (assumed equal to free Tt stream), 91 W airplane weight, lb U angle of attack of airplane center line, deg ratio of compressor-face pressure to standard NACA sea- 6 , P'

level pressure, -

a- Restriction/Classification t Cancelled CONF ......................... 0 . . 0 . .

0 . 0 . 0 .

...........

0 . 0 . . 0 . .

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

€ angle between airplane thrust axis and airplane center line, deg ratio of compressor-face temperature to NACA sea-level 0 , standard, Tt 518.4 Subscripts: b basic configuration S slats-locked-closed configuration AIRPLANE AND PROPULSION SYSTEM The test airplane is a swept-wing fighter airplane capable of super- sonic speeds. The 25-percent-wing chord line is swept back 45O, and the fuselage is characterized by a relatively flat bottom and positive cam- ber. Figure 1 presents a photograph of the test airplane, and figure 2 is a three-view drawing. Longitudinal control is effected by means of an all-movable stabilizer placed beneath the extended wing chord plane.

Each wing leading edge is equipped with an extensible slat consisting of five separate constant-chord sections and extending from about 25-percent to about 95-percent semispan. The slat is automatic in opera- Additional tion and opens as a function of local wing loading (ref. 1).

physical characteristics of the airplane are given in table I . Figure 3 presents the normal cross-sectional-area distribution in nondimensional form .

The propulsion system incorporates a sharp-lipped normal-shock nose inlet. The power plant is the 557 dual rotor turbojet engine with after- burner and a two-position iris-type nozzle. The bare engine military and afterburner thrusts are about 9,000 and l5,OOO pounds, respectively, at static sea-level conditions.

INSTRUMENTATION NASA recording instruments were installed in the airplane Standard to measure the following pertinent quantities: Airspeed Altitude Normal and longitudinal acceleration Angle of attack Restriction/Classification Cancelled 9 9 . 9 .. . 0 . 0.. ... ... .... ... .... 0..

Inlet-duct static and total pressure Jet-nozzle-exit total pressure Slat position Engine revolutions per minute, high speed and low speed Free-stream total temperature Stabilizer position A l l instruments were synchronized by a common timer.

Altitude and airspeed were determined by an NACA airspeed tube mounted on the nose boom, and angle of attack was measured by a vane attached to an arm projecting from the nose boom (airplane E of ref. 2 ) .

The vane was approximately 56.4 inches ahead of the inlet and 7 inches to the left of the center line of the boom.

THRUST AND DRAG DETERMINATION Gross thrust was determined in flight by measuring exit-nozzle total pressure and free-stream static pressure. Exit-nozzle total pressure was measured with an air-cooled cantilever-type probe inserted into the gas stream approximately in the jet-nozzle-exit plane. For most of the tests reported in this paper sonic flow was established at the jet-nozzle exit, permitting use of the following equation for gross thrust for both afterburning and nonafterburning operation The value of was determined from ground runs on a thrust stand and Cf was essentially the same for afterburning and nonafterburning operation.

Because Cf varies with exhaust-pressure ratio and because higher pres- sure ratios are attained in flight than on the ground, it was necessary to extrapolate the ground data to the higher pressure ratios. The extrap- olation was made in conformance with trends shown by altitude-chamber tests.

Figure 4 presents a typical thrust-stand calibration.

Duct total pressure was measured with three vertical rakes of four manifolded probes per rake each placed in the plane of the survey sta- The three rakes were connected to yield tion which was near the inlet.

one total pressure. Static pressure was measured with two connected wall static taps.

Ram drag was determined from the following equation: q-' Restriction/ .

C ONF D Classification Cancelled ......................... . . . . . . .

0 .

0 . 0 . . a * . . a

.......... ..f-'!;&?ENT$< 0. a

6 : : 0 . . a . . . . . .

..a . m . . a.. .......................

Local Mach number i n t h e duct was determined from measurements of t o t a l and s t a t i c pressure.

The derivation and l i m i t a t i o n s of t h e foregoing equations are d i s - cussed i n reference 3.

Following are t h e basic equations f o r the computation of the l i f t and drag c o e f f i c i e n t s : W

(an cos a + ax s i n a) - - F j s i n ( e + a)

CL = -

qs q s FJ Fr W

(&n s i n a - ax cos a) + - cos (E + a) - - CD = -

( 2 1

qs ss qs TESTS L i f t and drag were measured f o r both afterburning and nonafterburning operation. Most of the data a t Mach numbers g r e a t e r than about 0 . 9 were obtained w i t h t h e afterburner on, regardless of a l t i t u d e .

The f i r s t series of tests w a s performed with the leading-edge slats operating i n t h e normal manner. A Mach number range from about 0.63 t o about 1.44 w a s covered. For Mach numbers below about 1.1 the data were obtained a t a l t i t u d e s of about 20,000 f e e t , whereas t h e d a t a above about 1.1 were obtained a t a l t i t u d e s g r e a t e r than 20,000 feet and l a r g e l y a t 40,000 f e e t . The t e s t s consisted of push-down t u r n maneuvers (ref. 4 ) and speed runs.

The second t e s t s e r i e s w a s performed a t a l t i t u d e s from about 35,000 t o about 40,000 f e e t covering a Mach number range from about 0.87 t o 1.13.

For these t e s t s t h e various slat s e e e n t s were fastened t o t h e wing so t h a t no slat movement w a s possible under any f l i g h t conditions.

The t e s t s consisted of wind-up turns.

The f i n a l t e s t s e r i e s was i n s t i g a t e d as a check on the r e s u l t s of the f i r s t two t e s t s e r i e s and w a s obtained a t a later d a t e a f t e r the t h r u s t instrumentation had been removed. Therefore, it w a s possible Restriction/ CONFID Classification Cancelled

. . . . . . . ......................... - . - . .~

0 . 0 . 0 .

...........

. 0 . 0 . . . . 0 . : : 7

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

only to measure thrust-minus-drag of the airplane. Two flights were .

made at altitudes of 20,000 and 30,000 feet with the slats free-floating and sealed. The second flight was performed immediately after the first so that the ambient conditions at altitude were approximately equal.

The flights consisted of level-flight accelerations until the pilot felt that terminal Mach number was reached.

The test Reynolds number varied from about 25 million to about 48 million, based on the wing mean aerodynamic chord. Stabilizer posi- tion varied from Oo to about 18O, airplane nose-up.

ANALYSIS OF TKF: FINAL TEST SERIES Since the final tests were obtained some time after the initial tests, and since the thrust instrumentation had been removed from the airplane, it was possible to measure only thrust-minus-drag of the air- plane. However, the following analysis permits an estimate of any drag difference for the two configurations. For the speed runs employed in the final test, by assuming that the angles a and E are essentially 0, equation (2) can be simplified as follows: .

and Assuming

F j - Fr = Fn

Then At a given Mach number and duct efficiency the engine net thrust can be expressed as a function of the corrected engine speed, since Reynolds number and angle-of-attack effects are negligible.

Restriction/ CONFID Classification Cancelled ......................... . . . . . . .

a .

0 . 0 . . m a . . a 8 : : ..e 0 . 0 . . 8 . . ,**tObFe@ENT@f, : : : : .......... .......................

( 3 ) Therefore, a plot of the right-hand side of equation ( 3 ) against Mach .

number gives a direct indication of any drag difference between the two configurations for the speed runs.

ACCURACY The following accuracies are applicable f o r the results presented:

a, deg . . . . . . . . . . . . . . . . . . . . . . . . . +0.5 (average)

an, g . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

kO.05

a , , g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . kO.0025

F j , l b . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 150

Fr,lb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

M . . . . . . . . . . . . . . . . . kO.01 subsonically, supersonically

fO .02 transonically q (at M = 0.8 and $ = 37,500 ft), lb/sq ft . . . . . . . . . . .

?5

W,lb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ? l o o

A detailed discussion of the sources contributing to errors in neasurenent of these quantities is given in references 2 and 3 .

5 percent or less throughout the The error in lift coefficient is lift range presented. The accuracy of the drag coefficient depends Restriction/ Classification CONFID Cancelled 9 9 . 9 0 . . .........................

- - - - ..

. . 0 . 0 . . .

. . . 0 : : C!o&TDm*I:.. ........... i is

. 0 . 0 . ..e. . . . * .

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

primarily on the accuracies of thrust, angle of attack, longitudinal c acceleration, normal acceleration, weight, and Mach number. It is believed that the faired values of drag coefficient are accurate within +0.001 at low lift and higher values of dynamic pressure.

RESULTS AND DISCUSSION Figures 5 and 6 present lift coefficient plotted against angle of attack for the basic configuration and the slats-locked-closed configu- ration, respectively, for several Mach numbers; the data were obtained from the first and second test series. The Mach number variation for each plot was kept low; in the drag-rise region the variation was +0.01 and elsewhere was 50.02. Some nonlinearities are present in the curves; the nonlinearities at zero lift are due, possibly, to fuselage contribu- tion as a result of camber. For the slats-locked-closed configuration (fig. 6) the data are generally insufficient to determine any nonline- arities at low lift.

Figure 7 compares the lift data of figures 5 and 6 at three selected

Mach numbers. It will be noted that there are differences in the lift .

curves below the lift coefficient at which the slats start to open; the difference must either be the effect of the altitude change between the two sets of data, the result of some minor condition such as slat leak- During the final test series age, or the precision of measurement.

mentioned previously, the airplane was flown at a constant altitude with the slats free floating and locked closed and sealed; these tests showed conclusively that the effect was not the result of slat movement or leakage. However, in these tests only overall lift measurements were made, so the exact manner in which the lift is affected by altitude is not explained. (The airplane manufacturer's estimated data indicate changes in lift-curve slope due to structural flexibility equal to about one-half of those shown by these tests.) It should be noted that at a lift coefficient of about 0.3, about 0.3Omore angle of attack is required to produce a given lift coefficient at the low-altitude conditions than at the high-altitude conditions. Although the overall accuracy of measure- ment of angle of attack was only O.5O, it is believed that comparative measurements with the same system have a higher accuracy and hence the differences shown are real.

The slopes of the lift curves of figures 5 and 6 are plotted against Mach number in figure 8. Slopes were obtained for lift coefficients corresponding closely to 1 g flight for the test altitude and at a nominal weight of 22,000 pounds. At comparable Mach numbers the high- altitude data produced a higher lift-curve slope, with a maximum dif- 0.92, corresponding to a x ference occurring at a Mach number of about Restriction/ Classification CONFID Cancelled ......................... . . . . . . .

0 . 0 . 0 . ......

........ .C@&m&* : :* : : : lo: : 0 . 0 . .

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

Peak CLa of 0.068 for the high-altitude data. In addition, the high- altitude data show a more pronounced variation in the transonic region.

For the high-altitude data the lift-curve slope at a Mach number of 1.44 was approximately 62 percent of the value at a Mach number of 0.9.

Figure 9 presents drag coefficient plotted against lift coefficient for the basic configuration at altitudes of 20,000 and 40,000 feet.

The lift-coefficient and Mach number variation corresponds to the data of

figure 5. The drag-coefficient data corresponding to the lift curves

of figure 6 are not presented. The data were obtained at too low values of q, which produced larger errors in drag coefficient (i.e., for a a, etc.) than were encountered in the lower given error in thrust, altitude data and rendered presentation unsuitable.

The drag data from the final test series are presented in figure 1 0 as the drag-coefficient difference between the basic and slats-locked- closed configurations plotted as a function of Mach number at a given altitude. The analysis is given in the A N A L Y S I S O F THE F I N A L TEST SERIES section. The data indicate essentially negligible drag difference between the two configurations at altitudes of 20,000 and 30,000 feet for the test speeds; in addition, there was no difference between the terminal Mach numbers reached for the two configurations. The test lift coefficient was nominally 0.1 for both altitudes.

.)

Figure 1 1 shows the data of figure 9 plotted against Mach number for lift coefficients of 0 and 0.2. In the subsonic region the drag- coefficient levels remain relatively constant; in the supersonic region the drag-coefficient levels increase for Mach numbers greater than about 1-25. The drag-rise Mach number, taken as that corresponding to - - dCD - 0.1, was about 0.94 for both curves. At zero lift the supersonic dM drag-coefficient level was about three times as great as the subsonic drag level, which was about 0.01. Wind-tunnel data from reference 5 are presented for comparison. The data of reference 5 have been extrap- olated to the flight test value of Reynolds number using theoretical turbulent skin friction data. It should also be mentioned that geo- metric differences existed between the scale-model data of reference 5 and the full-scale airplane, notably in the tail thicknesses. In addi- tion, the model had no simulation of internal flow. Considering these differences, it is believed that reasonable agreement is shown between the two sets of data.

F i v e 12 presents the variation of ( L / D ) , , and CL for ( L / D ) , , with Mach number for the data of figure 9. The value of CL for ( L / D ) , , is confined between values of 0.25 and 0.35 for the test 1 Restriction/ Classification CONFI Cancelled .

. 0 . . 0 . .......................

0 . 0 .

: 0 . :11

: : to.: : c i N F e I A L . ........... 0 .

. 0 . 0 . .... 0 . .

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

range; the supersonic value of was about 35 percent of the ( L / D ) , , subsonic value.

Figure 13 presents the data of figures 5 and 9 plotted as

CD against CL2. Straight-line fairings were made for lift-coefficient- squared values less than about 0.16.

The slopes of the straight lines so obtained are a measure of the drag due to lift.

Figure 1 4 presents the variation of drag-due-to-lift factor with Mach number; the value increases through the Mach number range to a supersonic value about 137 percent the subsonic value. Altitude effects are assumed negligible.

The values of dCD / dCL2 corresponding to zero (1IcL,> and full

leading-edge suction (l/~tA) for low and high altitudes are also pre- sented; these data indicate a decrease in leading-edge suction as speeds are increased above subsonic values.

CONCLUSIONS Flight tests of the lift and drag of a swept-wing fighter airplane in the basic configuration and in the slats-locked-closed configuration for a Mach number range from 0.63 to 1 . 4 4 and altitudes of 20,000 feet to 40,000 feet led to the following conclusions: 1. At a nominal lift coefficient of 0.1 negligible drag-coefficient difference existed between the two configurations over a comparable Mach number and altitude range.

2. For the basic configuration at zero lift the supersonic drag level was about three times as great as the subsonic drag level, which was about 0.01, whereas the drag-due-to-lift factor increased about 137 percent over the test Mach number range.

3. At comparable Mach numbers the high-altitude data produced a larger lift-curve slope and showed a more pronounced variation of lift- curve slope in the transonic region than did the low-altitude data.

4. For the high-altitude data the lift-curve slope at a Mach number of 1.44 was approximately 62 percent of the value at a Mach number of 0.9.

High-speed Flight Station, National Aeronautics and Space Administration, Edwards, Calif., July 3, 1938.

Restriction/Classification Cancelled C 0°F ID .

1. Arabian, Donald D., Runckel, Jack F., and Reid, Charles F., Jr.: Aerodynamic Load Measurements and Opening Characteristics of Auto- matic Leading-Edge Slats on a 4 5 ' Sweptback Wing at Transonic Speeds.

NACA RM L53130, 1954.

2. Larson, Terry J., Stillwell, Wendell H., and Armistead, Katharine H.: Static-Pressure Error Calibrations for Nose-Boom Airspeed Installa- tions of 17 Airplanes. NACA RM H57A02, 1 9 7 .

3 . Beeler, De E., Bellman, Donald R., and Saltzman, Edwin J.: Flight Techniques for Determining Airplane Drag at High Mach Numbers.

NACA TN 3821, 1956.

4. Matranga, Gene J., and Armistead, Katharine H.: Flight Evaluation of the Effects of Leading-Edge-Slat Span on the Stability and Con- trol Characteristics of a Swept-Wing Fighter-Type Airplane During Accelerated Longitudinal Maneuvers at Transonic Speeds. NACA RM H58A03a, 1958.

5. Whitcomb, Charles F., and Lee, Edwin E., Jr.: Drag Investigation of a Swept-Wing Fighter-Airplane Model Incorporating Two Drag-Rise- Reducing Fuselage Revisions. NACA RM L55E24, 1955.

Restriction/Classification Cancelled CONFID . 0 . . 0 . .......................

...... 0 . 0 .

...... * & N F & & P ~ ........... 0 .

. 0 . 0 . .... 0 . . i i 13

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

TABLE I.- PHYSICAL CHARACTERISTICS OF AIRPLANE Wing :

A i r f o i l section . . . . . . . . . . . . . . . . . . . . NACA &A007

Total area (including ailerons and 83.84 sq f t

covered by fuselage), sq f t . . . . . . . . . . . . . . . . 385.21

Span, ft . . . . . . . . . . . . . . . . . . . . . . . . . . 38.58

Mean aerodynamic chord, f t . . . . . . . . . . . . . . . . . 11.16

Root chord, f t . . . . . . . . . . . . . . . . . . . . . . . 15.86

Tipchord, f t

. . . . . . . . . . . . . . . . . . . . . . . . 4.15

Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 0.262

Aspect r a t i o

. . . . . . . . . . . . . . . . . . . . . . . . 3.86

Sweep a t 0.25 chord l i n e , deg . . . . . . . . . . . . . . . .

0 Incidence,deg . . . . . . . . . . . . . . . . . . . . . . .

0 Dihedral, deg . . . . . . . . . . . . . . . . . . . . . . . .

0 Geometric t w i s t , deg . . . . . . . . . . . . . . . . . . . .

Aileron -

Area rearward of hinge l i n e (each), sq f t . . . . . . . . .

19.32 Span a t hinge l i n e (each), f t 7.81 . . . . . . . . . . . . . . .

Chord rearward of hinge l i n e , percent wing chord . . . . .

+15 Travel (each), deg . . . . . . . . . . . . . . . . . . . .

Leading-edge slat -

Span, equivalent, f t . . . . . . . . . . . . . . . . . . . 12.71

Segments . . . . . . . . . . . . . . . . . . . . . . . . .

. . .

Spanwise location, inboard end, percent wing semispan 23 - 3

Spanwise location, outboard end, percent wing semispan . . 89.2

Ratio of s l a t chord t o wing chord ( p a r a l l e l t o fuselage reference l i n e ) , percent 20 . . . . . . . . . . . .

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

Rotation, maximum, deg 15 Horizontal tail:

A i r f o i l section . . . . . . . . . . . . . . . . . . . . NACA 65A003.5

Total area (including 31.65 sq f t covered by fuselage),

s q f t . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 . 8 6

Span, f t . . . . . . . . . . . . . . . . . . . . . . . . . . 18.72

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

Mean aerodynamic chord, f t 5.83 . . . . . . . . . . . . . . . . . . . . . . .

Root chord, f t 8.14 Tipchord, f t

. . . . . . . . . . . . . . . . . . . . . . . . 2.46

Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . .

0.30 Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . .

3.54 Sweep a t 0.25 chord l i n e , deg . . . . . . . . . . . . . . . .

Dihedral, deg 0 . . . . . . . . . . . . . . . . . . . . . . . .

Travel, leading edge up, deg . . . . . . . . . . . . . . . .

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

Travel, leading edge down, deg 25 Control system , . . I r r e v e r s i b l e hydraulic boost and a r t i f i c i a l f e e l Restriction/Classification Cancelled CONFID

......................... e .e e . e

e . e .

1 4 1 i ........... ..t&fD~w&Q, : .... : :

e . e . . .... e . e . 0

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

TABLF, I.- PHYSICAL CHARACTERISTICS OF AIRPLANE . Concluded

Vertical tail :

Airfoil section . . . . . . . . . . . . . . . . . . . . . NACA 65A003.5

Area (excluding dorsal fin and area blanketed by

fuselage). sq ft . . . . . . . . . . . . . . . . . . . . . . . 42.7

Area blanketed by fuselage (area between fuselage contour line and line parallel to fuselage reference line through intersections of leading edge of vertical tail and fuselage

contour l i n e ) . sq ft . . . . . . . . . . . . . . . . . . . . . 2.45

Span (unblanketed). ft . . . . . . . . . . . . . . . . . . . . . 7 . 9 3

Mean aerodynamic chord. ft . . . . . . . . . . . . . . . . . . . 5.90

Root chord. ft . . . . . . . . . . . . . . . . . . . . . . . . . 8.28

Tip chord. ft . . . . . . . . . . . . . . . . . . . . . . . . . . 2.49

Taper ratio . . . . . . . . . . . . . . . . . . . . . . . . . . . 0.301

Aspect ratio . . . . . . . . . . . . . . . . . . . . . . . . . . 1 . 4 9

Sweep at 0.25 chord line. deg . . . . . . . . . . . . . . . . . . 45

Rudder -

Area. rearward of hinge line. sq ft . . . . . . . . . . . . . . 6.3

Span at hinge line. ft . . . . . . . . . . . . . . . . . . . . 3.33

Root chord. ft . . . . . . . . . . . . . . . . . . . . . . . . . 2.27

Tip chord. ft . . . . . . . . . . . . . . . . . . . . . . . . . 1.50

Travel. deg . . . . . . . . . . . . . . . . . . . . . . . . . . f20

.

Spanwise location. inboard end. percent vertical-tail span . . 3.1

Spamrise location. outboard end. percent vertical-tail span . . 44.8

Chord. percent vertical-tail chord . . . . . . . . . . . . . . 28.4

Balance . . . . . . . . . . . . . . . . . . . . . . . . . Aerodynamic

Fuselage :

Length (afterburner nozzle closed). ft . . . . . . . . . . . . . 45.64

Maximum width. ft . . . . . . . . . . . . . . . . . . . . . . . . 5.58

Maximum depth over canopy. ft . . . . . . . . . . . . . . . . . . 6.37

Side area (total). sq ft . . . . . . . . . . . . . . . . . . . 230.92

Fineness ratio (afterburner nozzle closed) . . . . . . . . . . . 7 . 8 6

Speed brake:

Surface area. sq ft . . . . . . . . . . . . . . . . . . . . . . . . 14.14

Maximum deflection. deg . . . . . . . . . . . . . . . . . . . . . 50

Power plant:

Turbojet engine . . . . . . . One Fratt & Whitney J57 with afterburner

Thrust (guarantee sea level). afterburner. lb . . . . . . . . . 1 5 . 000

Military. lb . . . . . . . . . . . . . . . . . . . . . . . . . . 9 . 000

Airplane weight. lb:

Basic (without fuel. oil. water. pilot) . . . . . . . . . . . . 2 0 . 262

Total (full fuel. oil. water. pilot) . . . . . . . . . . . . . . 2 5 . 400

Center-of-gravity location. percent c ' :

Total weight . gear down . . . . . . . . . . . . . . . . . . . . 30.2

Total weight . gear up . . . . . . . . . . . . . . . . . . . . . 30.2

Restriction/ Classification CONFI Cancelled • • • • • • 0 • o • 0 • 0 • • • : .

• • • • ..

• •• • 0 • o • .

tl 0 f-'

\Jl

.... •••• ••• ~

o • ••••• o o ••••• o ••••• • ••••• • ••••• • • •

• • ••••• ••••• • • o · : ° ·1· • • • • · •

E-2095 ....

, , position.

open in slats with airplane test the \ of Photograph - 1.

Figure Restriction/Classification Cancelled

o ~ H

J L

- 5 6 5 - 4

Figure 2.- Wee-view drawing of t h e t e s t airplane. A l l dimensions h inches.

Restriction/Classification Cancelled ..........

. 0 . . 0 . ...............

0 . 0 .

. 0.. : 37

: : :*.: : C O ~ I D ~ W . . ....... 0 . .

0 . 0 .

. 0 . 0 . ....

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

L I I I In - -

s

0 N O 0 0 9 9 9 Restriction/Classification Cancelled 3 " 1 . Q Q

cu

Restriction/ Classification CONFI Cancelled . 0 . . 0 . ....................

: : :*e: : COhD*ZAL::* ........... 0 . 0 . : 39

- - - - - - . . . . . . . .... 0 . . 0 . 0 .

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

k k t 3 V k E: Restriction/Classification CONFI Cancelled ......................... . 0 . . 0 . .

. . 0 . ......

2 6 i ........... . . *..'c~g*p*& : :e : : . . .

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

?

Restriction/ Classification Cancelled .

cu

-

a a , Lo r i V I a Q)

!

Io +J cd rl c n

*

cn k a ,

? I G

Q .

L I \o Restriction/Classification Cancelled CONF ......................... . 0 . . 0 . .

0 . e .

22: : ........... ..COBF~ENTfAfi : :a*: : : 0 . 0 . . . . . .... e 0 . 0 . .

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

O , -3 c,

r 0

, I I n i n?

I \ i

I I I I a , -0 L

\ \

\

\

\

I

a

(D Restriction/ Classification Cancelled CONJ?I .

.

k Restriction/Classification Cancelled CONF d rn C D c d .

F9 rn k a, 3 E:

- 2

rl l d k a ,

z

Lo k k +J d d .

W E : 4-10 ald 0 0 - P

1 8 4-1

d d ( - P o

Z V

d cd M d a al +J 0 c, rl PI -P c a , d .ri 4-1 4-1 al M cd

&

J I Restriction/ Classification Cancelled .

......................... 0 .

. 0 . . 0 . . 0 . 0 .

........

2.

...... csNFI~E&;Ac* ........... 0 . t t 2 5

. 0 . 0 . 0 . .

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

.

Restriction/ Classification Cancelled ......................... . 0 . . 0 . .

0 . 0 . 0 . ........

..........

0 . . . .

26 *e.: ....... .......................

h = 30,OoO f e e t M I I I t

- .04

.84 .88 .92 .96 I .oo I .04 I .08

M Figure 10.- Comparison of drag-coefficient difference between basic and slats-locked-closed configurations.

Restriction/Classification Cancelled CONFI .................

0 . 0 . 0 .

........

0 . 0 . : : 27

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

cu a E : cd k m 4J E: a J -4 d k k aJ (D d- 0 ' ' n Restriction/Classification Cancelled

I

!

h n \ -J v Restriction/Classification Cancelled CONFI QD

*

I c l V Restriction/ ~ Classification CONFI Cancelled ' 0 , (A i Restriction/Classification Cancelled .........................

. 0 . . 0 .

. 0 . 0 . 0 .

........ . 0 .

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

0 . 0 .

. . . . .

A l a I

\

i-

I

7- I

I

Restriction/ Classification Cancelled CONF NASA - Langley Fleld, Va.