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RM L57E17

RESEARCH MEMORANDUM

TIME-HISTORY DATA O F MANEUVERS PERFORMED BY A

REPUBLIC F-84F AIRPLANE DURING SQUADRON

OPERATIONAL TRAINING

By Harold A. Ham,er, John P. Mayer, and Donald B. Case

CON FkDENTIAL

NACA RM L57E17

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH MEMORANDUM

TIME-HISTORY DATA OF MANEUVERS PERFORMED BY A

REPUBLIC F-84F AIRPLCWE DURING SQUADRON

OPERATIONAL TRAINING

By Harold A. Hamer, John P. Mayer, and Donald B. Case

SUMMARY

Results of a control-motion study program with a jet fighter airplane are presented in time-history form. These time histories represent a cross section of tactical maneuvers performed by a Republic F-84F airplane during squadron operational training. Data were recorded at altitudes up t o about 42,000 feet, indicated airspeeds up to 600 knots, and at Mach numbers up to about 1.1. In addition, normal-load-factor data covering the entire flight investigation are separated into several different types of missions and summarized as envelopes of load factor plotted against indicated airspeed and Mach number.

INTRODUCTION

With a view toward providing information for the appraisal of the design requirements of future airplanes, the National Advisory Committee for Aeronautics has been conducting a statistical study of the pilot input, the corresponding airplane motions, and the loads imposed on a number of fighter-type airplanes during military training operations. The data obtained thus far have been swMlarized in references 1 to 3.

This paper presents data in time-history form of typical maneuvers performed by a Republic F-8b.F airplane during regularly scheduled operational training missions, which include the LABS (low-altitude bombing system) technique. “he time histories contain a cross section of the various tactical maneuvers and include the more severe maneuvers to show what maximum values of control and airplane motions can be expected during routine operational training. In addition, envelopes of normal load factor plotted against indicated airspeed and Mach number for several altitude ranges are presented for the complete flight program as well as for specific types of missions.

C 0 N F I D r n I A . L

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pressure a l t i tude, f t

indicated airspeed, knots

Mach number

aileron angle, deg

ai leron or spoi ler s t ick angle , deg

rudder angle, deg

s tab i la tor s t ick angle , deg

s t ab i l a to r angle, deg

a i le ron s t ick force , lb

rudder pedal force, lb

s t ab i l a to r s t i ck fo rce , lb

transverse load factor

normal load fac tor

longitudinal load factor

pitching velocity, radians/sec

pitching acceleration, radians/sec/sec

rolling velocity, radians/sec

rolling acceleration, radians/sec/sec

yawing velocity, radianslsec

yawing acceleration, radians/sec/sec

sideslip angle, deg

angle of attack, deg

mean aerodynamic chord, in . -

NACA RM L57E17 - 3

A standard U. S. Air Force F-84F-55RE airplane was used in t hese t e s t s . A photograph of the tes t a i rplane is shown i n f i g u r e 1. !be Republic F-84F is a single-place jet-propelled fighter-bomber airplane having a swept wing and empennage. The airplane is equipped t o c a r r y a wide assortment of weapons and ex terna l fue l tanks a t inboard and outboard stations along the wing.

A one-piece s tab i la tor i s used for longitudinal control. A damper is incorporated i n t h i s system t o decrease control sensit ivity and pre- vent overcontrol. In addition, w i t h the landing gear retracted, the control s t ick t ravel per degree of s tab i la tor movement i s increased t o reduce control sensit ivity. Included in the lateral-control system are spoilers which are located on the upper surface of both wings immediately forward of the f laps. The spoilers are actuated by the control-st ick la teral def lect ion. The spoilers operate when the landing gear i s i n t h e ret racted posi t ion and move in the up-direction only. As the a i leron moves up from 0' t o about l l o , the corresponding spoiler moves up propor- t iona l ly from 0' t o about 45O- and remains a t 45O f o r the remainder of a i leron t ravel . A l l control surfaces are hydraulically actuated. A hydraulically operated speed brake is ins ta l led on each s ide of the r ea r fuselage.

Neither the external appearance (except for the installation o f a nose boom) nor the weight and balance of the airplane was a l te red by the addition of the instrumentation. A three-view drawing of the F - 8 4 ~ test airplane is presented in f i gu re 2, i t s dimensions and physical character- i s t ics a re g iven in t ab le I, and i t s mass character is t ics are given in table 11. The moments of i n e r t i a given in t ab l e I1 fo r the various airplane configurations are values estimated i n accordance w i t h the best information available.

Standard NACA photographically recording instruments were used t o measure (1) the quantit ies defining the flight conditions - t h a t is, airspeed, altitude, Mach number, and speed-brake posit ion (open or closed), (2) the imposed control-st ick and control-surface motions along with the corresponding s t i c k and pedal forces, and (3) the response of the airplane i n terms of load factors, angular velocities, angular accelerations, angle of sideslip, and angle of attack. !be recorders were synchronized a t 1-second intervals by m e a n s of a common timing circuit . Microswitches incorporated in the a i rp lane bombing' system were used to i nd ica t e t he times at which the bomb switch and bonib release mechanism were actuated.

4 NACA l?M L57E17

In order to relieve the pilot of any recording-instrument switching pro- cedure and thus to assist in obtaining normal operation, a pressure switch was employed to operate the recording instruments automatically at take-off.

A standard two-cell pressure recorder connected to the airplane service system was used to measure the pressure altitude, indicated airspeed, and Mach number. The service system was of the usual total- pressure-tube and flush static-pressure-orifice type. (See fig. 2.)

The control-stick and control-surface deflections were recorded with a multiple-channel oscillograph having remote-recording variable- resistance-type transmitters installed at or near the control stick and control surfaces. Spoiler deflection was not measured at the surface since it varies directly with aileron deflection. The stick and pedal forces were recorded with the oscillograph by the use of NACA strain- gage-type transmitters mounted on the control stick and rudder pedals. A microswitch installed at the speed brake was used to indicate whether the speed brake was in the open or closed position.

Recorders containing gyroscopic sensing elements were used to record angular velocities and angular accelerations about three mutually perpen- dicular axes in which the longitudinal reference axis is the one commonly used for leveling the airplane. (See fig. 2. ) Load factors along these three axes were measured by an NACA magnetically damped, three-component recording accelerometer. The accelerometer was located in the vertical plane passing through the fuselage center line, 13 inches forward and I 2 inches above the average "in flight" center of gravity (21.4 percent of the wing mean aerodynamic chord and 3 inches below the fuselage reference line) .

The angle of attack and sideslip angle were measured by a flow- direction recorder in combination with vanes mounted on a boom extending from the upper right nose-gun port. (See figs. 1 and 2. )

ACCURACY

Table I11 is a summary of the -quantities measured and the estimated accuracies of the measurements. The accuracies are based on maximum possible instrument error plus the error resulting from preparing and reproducing the time-hist9ry figures. In order to give the correct response to the maneuver, all recording instrument elements were damped to about 0.63 of critical damping and their natural frequencies were selected to give the best compromise value which would minimize the magnitude of extraneous airplane vibrations.

In addition to the instrument and reproduction errors, it should be noted that no corrections are made to the recorded load factors for the effect of angular velocities and angular accelerations due to the displace- ment of the accelerometer from the airplane center of gravity. Since the accelerometer was near the center of gravity, these effects are negligible except for the effect of very large values of rolling acceleration on the recorded transverse load factor. The sideslip-angle and angle-of-attack measurements are uncorrected for angular velocity, sidewash, and upwash effects. It is estimated that the effects of sidewash and upwash increase the measured sideslip angle and angle of attack by approximately 5 percent and 10 percent, respectively. The effect of boom bending on the sideslip- angle and angle-of-attack measurements has been determined to be small.

TESTS

All the flights made during these tests were performed by service pilots undergoing regular squadron operational training. Measurements were obtained of 53 operational training flights with approximately 45 hours of flight time being recorded. The maneuvers performed during these tests include most of the tactical maneuvers that are within the capabilities of the airplane. The operational flights included such missions as LABS (low-altitude bombing system), dive-bombing, simulated strafing, in-flight refueling, formation flying, "rat-racing'' or dog fighting, and transition flying or familiarization (which usually consists of various acrobatic-type maneuvers). Data were recorded at altitudes from ground level to about 42,000 feet, at airspeeds varying from the stalling airspeed to about the maximum service limit airspeed (610 knots), and at Mach numbers up to about 1.1. Although not requested, most of the maneuvers were performed in smooth air except for those near the ground where turbulence was encountered.

Other than to request that the airplane be used in as many types of missions as were normally carried out by the squadron, no attempt was made to specify the type or severity of maneuvers. Since the pilots were aware of the instrumentation, it was stressed that this was not to restrict their normal handling of the airplane.

Sufficient film was available during each flight to obtain approximately 80 minutes of continuous records. This amount of time was generally enough to record the complete flight. The'flights were made with the clean airplane configuration and also with various external-store loading arrangements. Table IV is a swamasy of the flights from which time histories are presented.

6

METHOD AND RESULTS

NACA RM L’jT17

The sign convention used for the a i rplane motions and control posi- t ions is i l l u s t r a t e d i n figure 3. The basic results obtained are presented in figures 4 t o 64 as t i m e h i s tor ies of the measured quantit ies. The data presented herein are a small percentage of t h e t o t a l amount recorded during the tests, but are a cross section of a l l the operational maneuvers tha t were performed. In general, the maximum values of the control inputs and airplane responses obtained during the tests are included in these time-history figures.

In addition, envelopes of normal load factor plotted against indicated airspeed and Mach number for several altitude ranges are presented in f igures 65 t o 80. These plots contain normal-load-factor data, read a t 1/2-second intervals , for the complete flight program. The normal- load-factor data are separated into three types of missions, namely: (1) LABS (low-altitude bombing system) , (2) dive-bombing, and (3) general f lying, which includes a l l other types of missions.

The time h is tor ies shown in f i gu res 4 t o 64 were obtained by accu- rately tracing the f l ight-record traces (except for al t i tude, indicated airspeed, and Mach number, which were plot ted) direct ly on master sheets. The ver t ica l sca les on the f igures which conform to the instrument C a l i - brations are l inear and because of limited space are shown only i n one d i rec t ion for most quantit ies.

In some cases a time-history figure contains more than one maneuver. Care was taken t o begin and end each time-history figure a t a re la t ive ly normal steady-flight condition. The time-history figures are arranged according to t he maneuver c lass i f ica t ion shown in t ab l e V. Where a f ig - ure included more than one maneuver, i t s c lass i f ica t ion was established by that por t ion which was of greatest in terest . Included in the f igure legends are a description of the type of maneuver or maneuvers and the estimated in-flight airplane weight and center-of-gravity location, which were determined by interpolat ion with respect to total amount of f u e l used and the length of time for the flight including warm-up and taxiing. The flight from which each time-history figure was obtained i s a l so included so that reference can be made t o t a b l e IV.

In the t ime histories, the alt i tude is the NACA standard pressure a l t i t ude and the airspeed i s given i n terms of indicated airspeed which is defined as the reading of a differential-pressure airspeed indicator, cal ibrated in accordance with the accepted standard adiabatic formula t o indicate true airspeed for standard sea-level conditions only (uncorrected for instrument error). No corrections have been made to the indicated airspeed and Mach number for position error associated with the location

of the static orifices. Periods during which the speed brake was open are indicated on the f igures by dashed l ines and the words '%rake open.''

The control-surface and control-stick positions shown in the f igures were measured with respect to their neutral posi t ion. Only t h e l e f t ai leron posit ion was measured. Control forces were measured normal t o the s t i ck and pedals, and pedal force is given as the algebraic sum of the forces on both pedals.

The sideslip angle shown in the f igures i s the angle between the longitudinal axis and the projection of the re la t ive wind in the hor i - zontal plane of the airplane. The angle of a t tack i s the angle between the longitudinal axis and the projection of the re la t ive wind i n t h e ve r t i ca l plane of the airplane.

In the time-history figures, the angular velocities are shown as sol id t races and the angular accelerations as dashed traces. The time his tor ies of aileron st ick force are generally identified by traces with short dashes and those of rudder force with long dashes. Also, when the t race of any quantity intersects another, one of the t races i s dashed i n the region where t race ident i f icat ion might be confusing.

DISCUSSION

Although this invest igat ion w a s l imi ted to a re la t ive ly small amount of f l i g h t time, the data obtained represent a cross section of the maneuvers performed during operational training and include most of t he t ac t i ca l maneuvers within the capabili t ies of the F - 8 4 ~ airplane.

Time-History Figures ,

In general, the maximum values of the control inputs and airplane responses obtained during the tests are included in the t ime-history figures. Each group of time-history figures, as c lass i f ied in t ab le V according t o type of maneuver, will be briefly discussed:

Take-offs.- The time his tor ies shown in f i gu res 4 and 5 are repre- sentative values of control and airplane motions during take-off. In f igure 4, the unsymmetrical load of a fu l l 45O-gallon fuel tank (3,175 pounds) on the inboard pylon required that aileron angles up t o 10' be used t o maintain level wings at take-off. Also the automatic sh i f t i ng of the r a t i o of cont ro l - s t ick t rave l to s tab i la tor angle t o a higher value may be noted in f igure 4 (approximately a t times I 2 t o 20 seconds) after the landing gear i s retracted.

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8

It can be seen i n figure 5 that the longitudinal load factor during a jet-assisted take-off (JATO) i s increased from about 0 . 1 t o 0.3. As a matter of comparison the longitudinal load factor obtained during take- off for a normal weight ( f ig . 4) is about 0.2.

Turns.- Some of the highest normal load factors recorded during the tests were obtained in t u rns such as i n t h e l e f t and r igh t tu rns shown in f igure 7. In these turns, normal-load-factor values up t o about 6 were reached.

During the type of .maneuvering used i n r a t r a c i n g or t a i l chases, it was not unusual t o record long-duration turns made a t relatively high normal load factors. Figure 8 i s an example of such a maneuver.

Rolls .- In the ful l -a i leron rolls ( f igs . 14 and l5), ro l l i ng veloc- i t i e s up t o 3.8 radians per second were recorded; these values are about the m a x i m u m obtainable for the airplane for this f l ight condition. The maximum spoiler angle of 4 5 O i s reached during these rolls. The highest rolling accelerations recorded were ll radians per second per second a t indicated airspeeds of 350 knots t o 400 knots as shown in f igure 14 a t time 38 seconds and in f igure 15 a t time 30 seconds. It i s of i n t e re s t t o note that the highest roll ing accelerations were for the most par t obtained as the roll w a s being checked.

During rol l ing pul louts and ba r re l rolls (which are essent ia l ly rol l ing pul louts) , large rol l ing veloci t ies and roll ing accelerations were obtained i n combination with relatively high normal load factor a t high speed. (For example, see f ig . 18.) The magnitudes of yawing veloc- i t y , yawing acceleration, and sideslip angle, however, were small. According to calculations made for this type of maneuver being performed a t t h i s speed, t h i s r e s u l t might be expected. (For example, see ref . 2.) The calculations indicate that the values of these quantit ies would probably be considerably larger during rolling pullouts made a t lower speeds a t high l i f t coefficients. It can be seen i n f i g u r e 16 t h a t f o r about 1 minute a number of ba r r e l rolls were made during which time the normal load fac tor was continuously about 4.

Boost and speed-brake operation.- The time-history data s h a m i n figure 20 were obtained during the portion of a flight where the control- boost system w a s turned on a f t e r it had intent ional ly been turned off for a period of time. It is of interest to note the abrupt angular motions of the airplane together with relatively large angular accelerations as the controls are automatically trimmed when the boost i s restored.

The large aileron deflections resulted from the a i le ron s t ick force of about 12 pounds which the p i lo t was using when boost was restored. During most of the time when the boost was off , the a i leron s t ick force w a s off-scale (higher than 27 pounds). In another f l ight where the boost

was off, aileron-stick-force measurements made with a measuring element w i t h a higher range were found t o be as high as 80 pounds. In t h i s particular instance, the airplane carried an inboard 230-gallon external fuel tank which was almost fu l l (1,370 pounds). The 80-pound force was required to make a gradual bank t o 30' ( i n a direct ion away from the tank) a t 350 knots and a t an a l t i tude of 9,000 f ee t .

Figure 22 i s an example of the effect of speed-brake operation. For the par t icular f l ight condi t ion shown, a longitudinal deceleration of about O.5g was caused by use of the speed brake.

Dives.- Time h is tor ies of dives are presented in f i gu res 24 t o 29. Some of these dives pertain to what the p i lo t s termed "Mach runs." These were merely steep dives from high a l t i tude to speeds greater than a Mach number of 1. All high-speed dives shown in f i gu res 24 t o 29 were made with the clean airplane configuration, except f o r the one shown i n figure 29. For t h i s run the airplane carried a f u l l 230-gallon external fuel tank.

Upon examination of the Mach run time his tor ies ( for example, f i g . 24), some points of i n t e re s t can be noted such as the "heavy" tend- ency of t h e l e f t wing and the movement of the rudder t o angles up t o about 2' without any apparent intent of the p i lo t . (See the rudder force.) Also shown at or above Mach number 1 are the effects of shock waves on the sideslip-angle and angle-of-attack transmitter vanes. During these runs longitudinal decelerations up t o about 1.2g (fig. 26) were measured in the pul lout .

The highest dynamic pressure recorded during the tests was 1,200 pounds per square foot and was obtained i n a shallow dive t o law a l t i tude . (See f i g . 25 a t time 75 seconds.) This pressure occurred a t an a l t i tude of about 3,000 f e e t and a t a Mach number of about 0.95.

Dive-bombing runs.- In figures 30 t o 35 are shown typica l time h is tor ies of various types of dive-bombing maneuvers. The maneuver shown in f igure 30 i s tha t of an ordinary dive-bomb run which i s usually made from an a l t i tude of about 8,000 f e e t t o an a l t i t ude of about 3,000 or 4,000 f ee t . A glide-bomb run ( f ig . 31) i s made a t a smaller dive angle. The skip-bomb run shown in f igure 32 i s a maneuver made as c lose to the ground as possible, the bomb being skipped into the target. me runs shown in f igures 33 t o 35 are dive-bomb runs made a t higher al t i tudes. In these three runs practice bombs were dropped, the time a t which the pi lot re leased the bomb being indicated i n each f igure. In f igures 34 and 35 it can be noted that sideslip angles up t o 6' were obtained in lining up the dive-bomb runs..

LABS (low-altitude bombing system) runs. - The bombing runs shown i n figures 39 t o 46 are known as LABS maneuvers and are a technique developed

t o drop bombs from low a l t i tude . It can be seen from the f igures and by r e fe r r ing t o t ab l e IV t h a t some of these runs were simulated (no bombs dropped), some were made dropping small practice bombs, and some were made dropping a lY7OO-pound practice bomb. A LABS run consists of a straight-and-level, high-speed (approximately 300 knots) pass a t near- ground l eve l followed by an abrupt pull-up t o a prescribed normal load factor of 4, the bomb automatically releasing at a predetermined attitude angle. After the bomb i s released the half-loop is ca r r i ed t o completion, a 180° r o l l being made on top. It i s usually necessary to reduce the normal load factor as the speed decreases t o keep the a i rp l ane a t l i f t coe f f i c i en t s below the s ta l l and pitch-up region.

There are essentially two types of LABS missions, as indicated in the f igures . In an over-the-shoulder run the pull-up i s made d i r ec t ly ' over the t a rge t , the p i lo t in most cases putt ing the bo& switch to t he on-position a t the beginning of the pull-up. The bomb is r e l eased a t an airplane att i tude angle greater than go0, the angle depending upon such factors as a l t i t ude of target , wind velocity, and wind direct ion. In the cases where bombs were actually released, as indicated in the f igures, the re lease set t ings were l l 9 O for f igures 39 and 40 and 960 fo r f i g - ures 42 and 43. (5hose a t 96' are known as ver t ica l bomb drops.)

In the in i t ia l -poin t run, the pull-up i s made over what i s called an i n i t i a l p o i n t which i s selected ahead of the target . When the a i r - plane passes over the in i t ia l po in t , the p i lo t pu ts the b d switch to the on-position and then, after a predetermined elapsed time (depending on the distance from t h e i n i t i a l p o i n t t o t h e t a r g e t ) , makes the pull-up. I n comparison with the over-the-shoulder approach, the bo& r e l e a s e s a t a smaller at t i tude angle whiqe the airplane i s s t i l l a t a dis tance in f ront of the target . The angle set t ing for the ini t ia l -point LABS run shown in f i gu re 46 was 38.8'.

The high-speed portions of the runs often resulted in an extremely rough r ide (and occasionally caused t h e p i l o t ' s head t o h i t t h e canopy), especially when a 1,700-pound practice bomb was carr ied ( f igs . 41, 42, 43, 43, and 46). The turbulence encountered a t t h i s speed and a l t i t ude caused ai leron and rudder control-surface oscillations (about 1 cycle per second) with amplitudes as large as f-3' and +lo, respectively. It was learned that the pi lots were unaware of causing these motions and t h a t they were merely t ry ing to ho ld the s t ick f ixed and were res t ing the i r f e e t on the rudder pedals. During the longitudinal and direct ional osci l la t ions of the airplane, pitching accelerations as large as k2.3 radians per second per second, yawing accelerations as large as k1.4 radians per second per second, transverse load factors as large as k0.8, and normal-load-factor variations up t o kl.3 were measured a t the high airspeeds. These values of pitching and yawing acceleration and transverse load factor are higher than those measured i n a l l types of maneuvering during previous t e s t s made on other f ighter airplanes during squadron operations (ref. 1). Examination of the t ime histories

also shows that noticeable stabilator control-surface deflections occur with l i t t l e or no corresponding pilot-induced stick movement (under normal conditions the stabilator moves about 0.37' fo r lo of s t i c k t r a v e l ) . I n f a c t , a t times the stabilator and s t i ck a r e moving i n opposite directions. These inadvertent stabilator motions which are much more noticeable when the l,7OO-pound practice bomb i s on the a i r - plane cause the large incremental changes i n normal load factor. In addition, these inadvertent motions appear t o be associated with the severe directional oscil lations. It i s possible, therefore, that bending o r twisting of the fuselage-tai l combination could be enough to opera te the valve on the power control system and would force the s tabi la tor to move. (Calculations of t a i l loads and hinge moments indicated that the hinge moments were not high enough t o overpower the control system.)

In the bombing runs i n which a l,7OO-pound practice bomb was re.leased, an increase in normal load factor of about 1.5 (from 4 t o 5.5) was recorded i n an ini t ia l -point re lease ( f ig . 46) and about 0.7 (from 2.7 t o 3.4) i n an over-the-shoulder release (fig. 43).

Rat racing. - The term "rat racing" was used by the pi lots to def ine !

the type of maneuvering which may be employed in dog-fighting tactics. In f igures 47 t o 50 showing time h is tor ies of the data recorded during these rat-racing maneuvers, the tes t a i rplane was either being chased by or chasing another fighter airplane. Examination of the time h i s to r i e s shows tha t numerous la teral-control motions were used in t h i s t ype of maneuvering and that the occurrence of s t a l l s and pitch-ups was not unusual. In the pitch-up in figure-4g"it i s of spec ia l in te res t to no te where the airplane went completely QL& of mnt.ro_l_-(starting a t about time 93 seconds) i n a climb during a t a i l chase. It can be seen t h a t a sideslip angle o f a o and angles of attack as large as 26' (nose up) were recorded as well r p i t c h i n g and yawing veloci t ies of =ut 0-5 radian per second.

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Pitch-ups.- A n&tber of pitch-ups of varying degrees of severi ty were recorded during the tests, a few of which are presented i n f i g - ures 51 t o 58. Sometimes the pitch-ups were unintentional. In many cases they were intentionally performed since the service pilots did not consider the pitch-up of this a i rplane to be dangerous a t high a l t i - tude. Examples of intentional pitch-ups are given in figures 51, 55, 56, and 57. During the pitch-up shown i n figure 51, angles of a t tack up t o 27' (nose up), pitching velocities up t o 0.63 (nose up) radian per second, and pitching accelerations up t o 4.0 (nose down) radians per second per second were measured.

The pitch-up shown i n f i g u r e 57 was made at considerably higher indicated airspeed as compared with that for the other pitch-ups. Upon examination a t time 4.5 t o 6.5 seconds, it can be seen that dur ing this pitch-up the stabilator control-surface deflection does not correspond

t o the control-st ick motion. (Note tha t , for c la r i ty , the time scale has been expanded f o r this figure.) A n analysis w a s made of the maneuver by calculating the horizontal-tail loads and hinge moments from wind-tunnel data given in re fe rence 4. The results of th i s ana lys i s are given i n f igure 58 as time h i s to r i e s of the incremental structural (aerodynamic p lus iner t ia ) t a i l load and s t ructural h inge moment (taken about a hinge line inches forward of t h e t r a i l i n g edge a t the root chord). Included

i n t h e figure are the cont ro l s t ick and measured stabilator control- surface posit ions plotted on an enlarged vertical scale. In addition, the stabilator surface is shown as it would move under normal conditions.

4

The t a i l loads given in f i gu re 58 do not appear t o be excessive; however, the wind-tunnel t e s t s show tha t the hinge-moment-coefficient value increases rapidly w i t h angle of a t tack a t large angles of a t tack (above 12') due t o an outboard s h i f t of the t a i l load. These large hinge- moment coeff ic ients in combination with the re la t ive ly l a rge dynamic pres- sures produce hinge moments up t o about 14,000 foot-pounds, which is above the 9,000 foot-pound capacity of the power-control system. It can be seen i n figure 58 tha t the movement of the s tabi la tor surface i s apparently being affected by the large hinge moments which overpower the power control system.

Formation flying, ground control approach, in-fl ight refueling, and so for th . - Figures 59 t o 61 are presented t o show the type of control motions normally used while a p i l o t i s attempting t o remain i n a fixed l i n e of f l ight. For the instrument l e t d a m (ID) and ground control approach (GCA) shown i n f i g u r e 59 the t es t a i rp lane was f lying the wing posit ion with another airplane. For the formation flying shown i n f i g - ure 60 the t es t a i rp lane was pa r t of a large formation of airplanes. ?he time h is tor ies shown in f i gu re 61 are f o r an in-fl ight refueling operation i n which the p i lo t r epor t ed l i t t l e or no d i f f i cu l ty experienced i n hooking- up or remaining with the tanker airplane.

It can be seen that, i n the "break" or turn in the landing la rge ro l l ing ve loc i t ies (about 3 radians per second)

and large roll ing accelerations (about 9 radians per second per second) were obtained. These magnitudes, which can be considered unnecessary, are near the maximum ro l l ing capabi l i t ies of the airplane.

In figure 63 it is of in te res t to no te the dynamically unstable character is t ics of the airplane during the landing approach as sham by the large directional oscil lations. Figure 64 i s an example of a landing i n which a large amount of s ides l ip w a s used.

V-n and M-n Envelopes

All normal-load-factor data collected during the test program are s m a r i z e d as envelopes of load factor plotted against indicated airspeed and Mach number (herein denoted as V-n and M-n envelopes). These plots were made using automatic plotting equipment. These V-n and M-n envelopes are presented in figures 65 t o 80 along with the design and service diagrams fo r comparison of the tes t resul ts wi th the operat ional l imits . Essentially, the envelopes are composed of approximately 45 hours of load- factor data tabulated for every 1/2 second. In some of the f igures there are noticeable areas where the data are less dense ( for ins tance , in f ig . 65 between load factors of 1 and 3 and a t t he lower airspeeds up t o 400 knots). Because of the large mass of data in this region only enough points were plotted to adequately define the envelope.

Figures 65 t o 70 contain load-factor data, separated into three altitude ranges, for general flying which pe r t a in t o a l l t ypes of f lying other than that during flights scheduled solely for bombing missions. (The service limit load factor i s shown for the clean airplane configura- t ion since some of these data were obtained with this configuration.) Figures 71 t o 74 represent load factors obtained in two a l t i t ude ranges during LABS missions and figures 75 t o 78 include load factors recorded i n two a l t i t ude ranges during dive-bombing missions. (The service l imit load factor i s shown for the external stores configuration since a l l these data were obtained with this configuration.)

The boundaries of the combined load-factor data for general flying, LABS, and dive-bombing runs a t a l l a l t i tudes a re shown in f igures 79 and 80 plotted against indicated airspeed and Mach number. It i s indicated that, as i n similar t e s t s made with other f ighter airplanes (ref. 1) , the service pi lots ut i l ized the posi t ive port ion of the V-n or M-n diagram but did not approach the negative-load-factor limits.

Inspection of figures 65 and 66 shows that the data do not extend up t o t h e s ta l l boundary of the V-n diagram whereas they do when plotted on the M-n diagrams. This condition can also be noticed for some of the other sets of V-n and M-n plots . This condition occurs because the effect of a l t i t ude on the M-n s ta l l boundary i s large as compared with that on the V-n s ta l lboundary. The data points in f igure 66 which f a l l on or near the M-n 5,000-foot-altitude s ta l l boundary were obtained near sea level . I n the M-n diagram the sea-level s ta l l boundary w i l l be higher than the 5,000-foot s ta l l boundary whereas the sea-level and 5,000-foot boundaries i n t h e V-n diagram are about the same; therefore, the position of the data with respect to the sea-level boundary w i l l be the same i n t h e two figures.

a In the envelopes for general f lying for the alt i tude range from sea

l e v e l t o 15,000 f ee t ( f i g s . 63 and 66), the highest load factors of 6.4 and 6.1 were measured during the pullout from the simulated dive-bomb run -

shown in f igure 27. The other high load factors above 6 were recorded during the tight turns shown in f igure 7. For the general-flying maneuvers a t 15,000 feet t o 30,000 f ee t ( f i g s . 67 and 68) , the load fac tors shown a t 5.5 and above were recorded in the pitch-up shown in f i gu re 57. The load factors above the s ta l l boundary in t he v i c in i ty of 275 knots were, of course, also obtained during pitch-ups. The negative-load factors shown in f i gu res 67 and 68 were recorded during portions of f l i g h t s where load factors of less than 1 were intent ional ly appl ied in checking out new engines. In the envelopes for 30,000 f e e t t o 42,000 f ee t , a l l load factors above 2.8, as well as those above t h e s t a l l boundary ( see f ig . 69) a t the lower airspeeds, resulted from pitch-ups.

CONCLUDING REMARKS

A f l igh t inves t iga t ion conducted on a Republic F - 8 4 ~ service airplane during routine squadron operations has indicated that the service pi lots ut i l ized the posi t ive port ion of the V-n or M-n diagram when per- forming various t a c t i c a l maneuvers. In the f l igh t time recorded in these t e s t s , however, the negative normal-load-factor limits were not approached.

The maximum values of roll ing velocity and rol l ing accelerat ion recorded were 3.8 radians per second and 11 radians per second per second, respectively. The m a x i m u m values of pitching velocity and pitching acceleration resulted from pitch-ups and were 3.8 radians per second and 4 radians per second per second, respectively. In addition, directional osci l la t ions a t the high speeds during LABS (luw-alt i tude boding system) missions produced yawing accelerations up t o 1 .4 radians per second per second and transverse load factors up t o 0.8. All these values are considerably higher than any previously recorded during similar tests made with other f ighter airplanes.

Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Langley Field, Va., April 19, 1957.

I , NACA RM L57E17 - 15

REFERENCES

1. Mayer, John P., Hamer, Harold A., and HUSS, Carl R. : A Study of the Use of Controls and the Resulting Airplane Response During Service Training Operations of Four Jet Fighter Airplanes. NACA RM ~ 3 3 ~ 2 8 ,

I Il f 1954 '

i 2. Mayer, John P., and Hamr, Harold A.: A Study of Service-Imposed, Maneuvers of Four Jet Fighter Airplanes i n Relation t o Their Handling Qualities and Calculated Dynamic Characteristics. NACA RM L55E19, I' 1 1955.

I 3. Mayer, John P., and Hamer, Harold A.: Applications of Power.Spectra1 Analysis bkthods t o Maneuver Loads Obtained on Je t Fighter Airplanes During Service Operations. NACA RM L56J15, 1957.

4. Hallissy, Joseph M., Jr., and Kudlacik, Louis: A Transonic Wind-Tunnel Investigation of Store and Horizontal-Tail Loads and Some Effects of Fuselage-Af'terbody Modifications on a Swept-Wing Fighter Airplane. NACA RM ~ 5 6 ~ 2 6 , 1956.

16 . NACA FM L5P17

TABU3 I . . DIMENSIONS AND PHYSICAL CHARACTERISTICS

OF THE REPUBLCC F - 8 4 ~ TEST AIRPLANE

Wing ( t rue dimensions and areas): Total wing area (including flaps. ailerons. and 50.6 sq f t

covered by fuselage). sq f t . . . . . . . . . . . . . . . . Span. f t . . . . . . . . . . . . . . . . . . . . . . . . . . Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . Taper ratio. (Tip chord/Root chord) . . . . . . . . . . . . .

l i n e ) . i n . . . . . . . . . . . . . . . . . . . . . . . . . l i n e ) . i n . . . . . . . . . . . . . . . . . . . . . . . . . to fuse lage cen ter l ine) . in . . . . . . . . . . . . . . . . . . chord (para l le l to fuse lage cen ter l ine) . in . . . . . . . . .

Theoretical root chord (measured paral le l to fuselage center

Theoret ical t ip chord (measured paral le l to fuselage center

Mean aerodynamic chord (at wing s ta t ion 90.6 measured normal

Distance from nose to leading edge of wing mean aerodynamic

Vertical location of wing mean aerodynamic chord (25 percent) in re fe rence to fuse lage cen ter l ine . in . . . . . . . . .

Sweepback (measured a t 25-percent-chord l i ne ) . deg . . e . . Theoretical root and t i p a i r f o i l s e c t i o n (measured normal t o

Airfoi l th ickness ra t io (measured para l le l to fuse lage 25-percent-chord l i ne ) . . . . . . . . . . . . . . . . UACA

center line). percent chord . . . . . . . . . . . . . . . . Incidence. deg . . . . . . . . . . . . . . . . . . . . . . . Geometric twist. deg . . . . . . . . . . . . . . . . . . . . Cathedral. deg . . . . . . . . . . . . . . . . . . . . . . . Aileron area (two). sq f t . . . . . . . . . . . . . . . . . . Aileron s t a t i c con t ro l limits. deg . . . . . . . . . . . . . Spoiler area (two). sq f t . . . . . . . . . . . . . . . . . . Spoiler s ta t ic control limit (up only). deg . . . . . . . . . Total area. sq f t . . . . . . . . . . . . . . . . . . . . . . Span. f t . . . . . . . . . . . . . . . . . . . . . . . . . . Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . Taper r a t i o . . . . . . . . . . . . . . . . . . . . . . . . .

Stabilator :

Chord (measured paral le l to fuselage center l ine) . in . . . . Wan aerodynamic chord. i n . . . . . . . . . . . . . . . . . Root and t i p a i r f o i l s e c t i o n (measured normal to leading

Airfoi l th ickness ra t io (measured para l le l to fuse lage edge) . . . . . . . . . . . . . . . . . . . . . . . . . NACA

center line). percent chord . . . . . . . . . . . . . . . . Sweepback. deg . . . . . . . . . . . . . . . . . . . . . . . Dihedral. deg . . . . . . . . . . . . . . . . . . . . . . . Tail length (25 percent wing mean aerodynamic chord t o

25 percent stabilator mean aerodynamic chord). f t . . . . .

324.7 33.6 3.45

0 578

148.6

86.0

EO .5

187.6

14.44 40

64AOlO

8.1 1.5

0 3-5

27.9 k18 7.7 45

3.59 1.0

48.0 48.0

64AOO9

6.9 40 0

19.6

NACA RM L5i7317 . 17

TABIX I . . DIMENSIONS AND PKYSICAL CHARACTERISTICS

OF 'IIRE REPUBLIC ~ - 8 4 F TEST AIRPLANE . Concluded

Stat ic control limits. deg from neutral position Wailing edge up (gear up) . . . . . . . . . . . . . . . . . 11.25 Trailing edge up (gear down) . . . . . . . . . . . . . . . . 16.5 Trail ing edge down (gear up) . . . . . . . . . . . . . . . . 4.25 Wail ing edge down (gear down) . . . . . . . . . . . . . . . 4.5

area). sq f t . . . . . . . . . . . . . . . . . . . . . . . . Vertical tai l :

Total area (does not include dorsal and ven t r a l f i n

Theoretical root chord (measured paral le l to fuselage center

Location of theoret ical root chord above fuselage center

Dorsal f i n area. sq f t . . . . . . . . . . . . . . . . . . . 2.02 Ventral f i n area. sq f t . . . . . . . . . . . . . . . . . . . 2.12

Aspect r a t i o . . . . . . . . . . . . . . . . . . . . . . . . . 1.60

Mean aerodynamic chord (located 60 inches above fuselage

Root and t i p a i r f o i l s e c t i o n (measured normal t o 25-percent-

l i n e ) . i n . . . . . . . . . . . . . . . . . . . . . . . . . 87.0

l ine . i n . . . . . . . . . . . . . . . . . . . . . . . . . . 15.25

Span (from theoretical root chord). f t . . . . . . . . . . . . 8.5

Tape r ra t io . . . . . . . . . . . . . . . . . . . . . . . . . 0.402

center l ine) . in . . . . . . . . . . . . . . . . . . . . . . 64.6

chord l ine) . . . . . . . . . . . . . . . . . . . . . . NACA 6 4 A O l l Sweepback (measured a t 25-percent-chord l ine) . deg . . . . . . 41.27

Rudder area. sq f t . . . . . . . . . . . . . . . . . . . . . . 10.37 Rudder s ta t ic cont ro l . l imi t s , deg . . . . . . . . . . . . . . k23.5

Length (excluding nose boom). f t . . . . . . . . . . . . . . . 38.42 Maximum width. f t . . . . . . . . . . . . . . . . . . . . . . 4.17 Frontal area (excluding canopy). sq f t . . . . . . . . . . . . 19.43 Speed brake area ( two) . sq f t . . . . . . . . . . . . . . . . 11.7

Power plant (one) . . . . . . . . . . . . . . . . . . . . Wr.J-65-B3 Airplane serial n-er . . . . . . . . . . . . . . . . . . AF 52-6984

Tail length (25 percent wing mean aerodynamic chord t o 25 percent vertical-tail mean aerodynamic chord). f t . . . . 18.0

.. Fuselage :

General :

Weight of empty 230-gallon tank. l b Inboard . . . . . . . . . . . . . . . . . . . . . . . . . . 173 Outboard . . . . . . . . . . . . . . . . . . . . . . . . . . 90

Weight of outboard pylon (one). lb . . . . . . . . . . . . . 80 Weight of 1 gallon of f u e l (assumed) . l b . . . . . . . . . . . 6.5 Spanwise location of inboard pylon. i n . . . . . . . . . . . . 44

Weight of empty 45O-gallon tank. l b . . . . . . . . . . . . . 249

Spanwise location of outboard pylon. i n . . . . . . . . . . . 149.5

18 II$ NACA RM L57E17

!CABLE 11.- MASS CHARACTERISTICS OF THE REPUBLIC F-&F TEST -LANE

I I Measured

Clean; f u l l f u e l

Clean plus one empty 23O-gallon tank (inboard); f u l l i n t e r n a l fuel

Clean plus one fu l l 230-gallon tank (inboard) ; fu l l i n t e r n d f u e l

Clean plus two empty 230-gallon tanks (inbogrd); f u l l in te rna l f u e l

Clean plus two fu l l 230-gallon tanks (inboard) j f u l l i n t e r n a l f u e l

18,540 20.0

18,710 20.2

20,210 20.3

18,890 20.4

21,880 20.6

Clean plus one empty 4w-gallon

Clean plus one fu l l 4W-gallon

f u e l

Clean plus one empty 4w-gallon tank (inboard) and one 1,700-pound pract ice bomb (inboard); f u l l i n t e rna l fue l

20,490 18.1

Clean plus one fu l l 450-gallon tank (inboard) and one J.,7OO-pound pract ice bomb (inboard); fu l l in t e rna l fue l

Clean plus one empty 450-ganon

23,410 17-9

tank (inboard), one I 1,700-pound practice bomb (inboard), and two empty I 20,670 I 18.4 230-gallon tanks (outboard) ; I

Clean plus one f u l l ky-gallon tank (inboard), one l,7OO-pound practice bomb (inboard) , and two fu l l 230-gallon tanks (outboard) 5 fu l l i n t e rna l fue l

18,000 32,900 48,400

20,200 34, 600 50,400

19,300 50,000 34,600

21,600 53 100 37,400

2'3,300 51,100 3 4 , s O O

38,200 70,200 40, 500

%lean includes two inboard and two outboard pylons. %eight of 200-pound p i l o t included. Ecternal fuel i s normally used before internal fuel.

NACA RM L5m17

TABLE 111.- SUMMARY OF MEASURED QUANTITIES

Quantity Maximum instrument error plus

reproduction error

Pressure altitude. Hp. f t . . . . . . . . . . . . . . . . . . *+200

Indicated airspeed. Vi. knots . . . . . . . . . . . . . . . . *+lo Mach number. M . . . . . . . . . . . . . . . . . . . . . . . *. 0 . 02 Stabilator angle. ... deg . . . . . . . . . . . . . . . . . . rto. 2 Rudder angle. t+. deg . . . . . . . . . . . . . . . . . . . . 20.5 Aileron angle. tja. deg . . . . . . . . . . . . . . . . . . . 20.5 Spoiler angle. deg . . . . . . . . . . . . . . . . . . . . . 22.0 Stabi la tor s t ick angle. 6ss. deg . . . . . . . . . . . . . . 20.5 Aileron or spoiler stick angle. deg . . . . . . . . . . Stabi la tor s t ick force. F.. lb . . . . . . . . . . . . . . . Aileron s t ick force. Fa. lb

For flights 2 through 25 . . . . . . . . . . . . . . . . . For flights 26 through 53 . . . . . . . . . . . . . . . . .

Rudder ...... Fr. lb . . . . . . . . . . . . . . . . . . . . Normal load factor. nV . . . . . . . . . . . . . . . . . . . Transverse load factor. nT . . . . . . . . . . . . . . . . . Longitudinal load factor. nL . . . . . . . . . . . . . . . . Pitching velocity. q. radians/sec . . . . . . . . . . . . . . Yawing velocity. r. radianslsec . . . . . . . . . . . . . . . Rblling velocity. p. radians/sec . . . . . . . . . . . . . . Pitching acceleration. 6. radians/sec Yawing acceleration. i., radians/sec 2

Rolling acceleration. e. radians/sec 2

2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

Sideslip angle. p. deg . . . . . . . . . . . . . . . . . . . Angle of attack. a. deg . . . . . . . . . . . . . . . . . . . Time. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . *Value does not include position error .

+o .5 20.8

+o .8 23.0 25.0 20.1

50 . 05 fO .05 50.02 10.02 10 . 15 *o .2

20.08 20.3

*+O .4 *Af0.4

20.2

.

20 NACA RM L57E17

Configuration

( 3

I 3 I B I 18,250 I 20.3 I Clean (minus one outboard pylon) I Wansition

I 7 I D I 18.150 I 20.1 t Clean (minus ivo outboard wlonsl I Wansition

9 E 18,200 20.0 Clean (minus two outboard pylons) Wansition

l l F 17,850 20.8 Clean (minus tvo outboard pylons) Wansition

I3 0 18,150 20.1 Clean (minus tvo outboard pylons) Wanaition

14 H 18,200 20.0 Clean (minus tvo outboard pylons) Rat-racing and dive-bmbing

15 I G I 18,250 I 19.9 I Clean (minus tvo outboard pylom) Transition

16 E 19,800 19.7 23O-g&on tank on right inboard pylon Bambing, strafing, and rat-racing (minus two outboard pylons)

17 C 19,800 19.7 23O-gallm tank on right inboard pylon Farmation flying lminus tvo outboard uvlcma)

19 J 20,000 20.1 2jo-g&M tank on right inboard pylon TransitiOn

20 J 20,000 20.1 230-gallon tank on right inboard pylon Waneition

25 K 19,900 20.0 230-gallon tank on right inboard pylm Farmation flying

26 I L I 18,300 [ 20.5 [ Clean Wansition

l,7OO-pound practice bomb on l e f t inboard pylm, and 3 - p W practice

aAirplane equipped v i t h tvo outboard pylons ("nless otherwise stated) snd vith two inboard p y l m . bWansitim refers to familiarization flylng vhich usually includes various acrobatic-type maneuvers.

21

TABU V.- MANEWER CLASSIFICATION

Type of maneuver: Take-offs . . . . . . . . . . . . . . . . . . . . . . Turns . . . . . . . . . . . . . . . . . . . . . . . . Rolls . . . . . . . . . . . . . . . . . . . . . . . . Immelman, loop, rudder kick, and so fo r th . . . . . . Dives . . . . . . . . . . . . . . . . . . . . . . . . Strafing runs . . . . . . . . . . . . . . . . . . . . LABS (low-altitude bombing system) runs . . . . . . . R a t -rac ing . . . . . . . . . . . . . . . . . . . . . Pitch-Ups . . . . . . . . . . . . . . . . . . . . . . Formation flying, ground control approach, in-f l ight

Dive-boding runs . . . . . . . . . . . . . . . . . .

refueling, and so fo r th . . . . . . . . . . . . . . Landings . . . . . . . . . . . . . . . . . . . . . .

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

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

. . . . . .

Figures 4 t o 5 6 t o 9

10 t o 18 19 t o 23 24 t o 29 30 t o 35 36 t o 38 39 t o 46 47 t o 50 51 t o 58

59 t o 61 62 t o 64

la ..

Figure 1.- Republic F-84F t e s t airplane. L-91212

NACA RM L57E17

head

r Service-system s t a t i c - pressure orifices

F

reference line

z r I Fuselage

reference line

Figure 2.- Three-view drawing of Republic F-&F tes t a i rp lane .

24 NACA RM L 5 p l 7

RE

Figure 3 . - Body-axis system showing posit ive directions.

Ressure

Lodd factor

L

Angular i- 1-

rad /sec velocity,

Angle of attack, de9

load factor

-i I Angular

accet. rad/sec 2

Sideslip angle, de9

Figure 4.- Data obtained during take-off, unsymmetrical loading. A i r - plane weight, 21,450 pounds; center of gravity at 19.0 percent 5; f l i g h t 47.

26 NACA RM ~ 3 7 ~ 1 7

R e S s V e altitude, f t

lndicuted airspeed, knots

Can i, ol surface position,

i Rudder

Load r factor

L

Rudder

Lt " € - " - I I I I I

I I I -l 0 Ailwon

14 control Lt. I

' 3 - I I I I I 12 position, stick

Rear ".A I I 0 Stabilator v

3 0 R t Aileron Jk

Angular veloclty, rad /sec

I L

Angle of attack, deg

Normal load factor

1 J Angular

accet. rad/sec2

Sideslip angle, deg

Figure 5 . Data obtained during take-off ( J N O ) . Airplane weight , 27,050 pounds; center of gravity at 24.8 percent E ; f l i g h t 49.

I 111 I I

I

I i

NACA RM ~57~17 - Ressure altitude, f t

lndlcated airspeed, knots

Mach number

subface position, deg

I

L Rudder

Load r factor

I L

r

12 Right On

40

0

0 Ailem 7 stick position,

Rear deg

U P W position, spoiler

aileron 0 Left

Rudder 0 Lt, Lt 14 Control

IO 12 8

TE. UP 0 s+abilatwJ

Stabilator 0 100 R t.

force, Ib 0

I 2

0 Aileron 7 Rt.

Stick

2 o Stablator Push Ib, R t Transverse 0 12

5 2

Long!?$ I

Pitch 0 UP

Dn.

2

0 Pitch

-2

UP

Dn.

I Angular

velocity, rad /sec

L Angle of

attack, deg

Time, sec

3

Nwrnal load factor

1 J Angular occet. rodkc'

Sideslip angle, de9

Figure 6. - Data obtained f o r a low-speed turn, f laps down. Airplane weight, 18,100 pounds; center of gravity at 21.0 percent e; f l i g h t 2.

Pressure altitude, f t

Indicated airspeed, knots

Control r surface position,

15,000

10,000

5,000 600 500 400 Mach 300 number 200 IO0

Normal

0 IO 20 30 40 50 60 Time, sec

Figure 7. - Data obtained during tight turns. Airplane weight, 15,530 pounds; center of gravi ty a t 22.3 percent 5 ; f l i gh t 14.

NACA RM L5n17 e 29

Mach number

l0,OOO

Pressure altitude, f t 5,000

0 600 1.2 500 I .o

airspeed, 400 .8 knots 300 .6

Indicated

2 0 0 .4 IO0 .2

12

r Right

Dn Left UP only pasimn, olleron 0

Control Rudder 0 surface position, Lt

IO 8

40 spoiler

0

0 Aileron 7 14 Control

stick

Lt,

deL 12 position, 7: E. Rear de9

UP 0 StobilotorA Stobilator 0

100 Rt

Rudder force, Ib 0 2 r Transverse 0

2

L Long. g

12

0 Aileron 7 0 Stabilotor Ib

Rt.

Stick force,

R t Push 12 _1

Load 5 Normol

I factor

2 UP

0 Pitch

factor load

Fwd.

UP r Dn.

3 R t. Rt. 1 velocity, Roll 0 0 Roll Angular

rad /sec Lt. Lt, accet.

.5 .8

Pitch 0 Dn.

75 -2

Angular 5

-3 -5 rod/sec2

Yaw Rt.0 Lt. Lt. _1 0 Yaw

6 -.8

R t.

-.5

Nose It. Sideslip

Nose rt. deg 0 angle,

10 Angle of Nose 6

attack, deg

UP 0

0 10 20 30 40 50 60 Time, sec

(a>

Figure 8.- Data obtained during series of turns. Airplane weight 17,400 pounds; center of g rav i ty a t 22.5 percent 5; f l i g h t 18.

NACA RM L57E17

Pmsure altitude, ft I 0 " W 5,000


surface position, deg

Rudder

r Laad

. factor

L

r

- - 0

600 500 400 1 Vi

300- M

- - -

-

200 1 IO0

12 r Dn E

alleron 0 Left

Rudder O Y - ?

sa

6, - *"

Lt z IO - -

force, Ib 0

Transverse 0

Long.

Pitch 0 UP

Dn.

Angular I

velocity, rad /sec

1 Angle of

attack, deg

I

i I .2 I .o .8 .6 .4 .2

Mach number

Right 40 spoiler

0


12

UP only w n , deg

Lt.

stick position,

Rear deg o Stabilator J 12

0 Aileron 1 Rt.

Stick

Yaw

I

Normal

factai load

1 Angular

accel. rad/sec2

J .:

Sideslip angle, deg

Time, sec

(b)

Figure 8.- Concluded.

Ressure altitude, f t


r Control

surface position, de9

400 300 200 I O 0

c

Mach number

12 Right

On 40

UP only position, spoilw

aileron 0 Left 0

Rudder 0 Lt i" Control stick position,

0 Ailerm

Lt 14 10 8 12

TE. Rear deg

UP 0 S tab i l a to r i Stabilator 0 48

100 0 Aileron 1 2 OpushStobilator ,b

Rt

Rudder force, Ib 0 Rt Stick

force,

R t r O 12

5

I

2

0 Pitch

Load factor 2

Longfwdg

UP Pitch 0 [ Dn.

UP

Dn. -.5 -2

3 5 R t. Rt. L t.

-5

Angular velocity, Roll 0 0 Roll rad /sec -3

Lt.

Yaw Rt.0

.5 .a 2. 0 Yaw

R t.

Lt. Lt. -.5 -.8

Angle of ottack, de9

Nos;om 1 1 1 T'""' ! It.

UP 0

0 10 20 30 40 50 60 Time, sec.

_I

Normal load factor

1 1 Angular

accel. rad/sec2

Sideslip angle, deg

Figure 9.- Data obtained during a tight turn. Airplane weight, 17,500 pounds; center of grav i ty a t 22.2 percent E ; f l i g h t 34.

NACA m ~ 5 7 ~ 1 7


20,000

Ressure altitude, f t 15,000

10,000 600 I .2 500 I .o 400 .e Mach 300 .6 number 2 0 0 .4

r IO0 .2

12 Right

Dn Left UP only p m n , aileron 0

Rudder 0

IO 8

40 spoiler

0

0 Aileron

surface position, Lt 14

Lt. i" Control stick

12 position,

TE. Rear deg

UP 0 S tab i l a to r l Stabilator 0

100 R t.

Rudder force, Ib 0

Rt 12 Ibl

12

0 Aileron 7 Rt.

Stick

2 OpushStabilator force'

r Transverse o L a i d

factor

L

Angular l- L

rad /sec velocity,


I I I I I I I 0 I O 20 30 40 50 60

Time, sec

2

Normal load factor

1 1 Angular accet. rad/sec2

Sideslip angle, deg

Figure 10. - Data obtained for a loop followed by roll ing pullout. A i r - plane weight, 16,750 pounds; center of gravity at 22.1 percent E ; f l i g h t 3.



r Control

surface position,

15.000 - I - - -

1 0 . 0 0 0 ” ~ Hp

Figure 11.- Data obtained during eight consecutive barrel rolls. A i r - plane weight, 16,300 pounds; center of grav i ty a t 22.0 percent E ; f l i g h t 7.

34



r coitrol

surface position,

der

300 200 IO0

L Angle of

attack, de9

.5

Yaw 0

-.5

R t.

Lt.

IO Nose

UP 0

I I I 0 IO

I

2o Time, sec. 30

I

1.2 I .o .8 .6 .4 .2

Mach number

-0 dea

12

0 Stab i l a twd 12

position,

Rear

Stick

Nwmal

factor load

2

oUp Pitch

-2 5

0 Roll

-5

Dn.

R t.

Lt.

.+]:Nose rt.

Nose It.

) 50 60

1 Angular

accet. rad/sec2

J Sideslip angle, de9

Figure 12.- Data obtained during r o l l s . Airplane weight, 16,900 pounds; center of g rav i ty a t 22.6 percent 5; f l i g h t 18.



35,000

30,000

25,000 600 1.2

Indicated 5 0 0 1 .o airspeed, 400 knots 300 .6

.8 Mach number

200 .4 IO0 .2

12 Right

r Dn 40

UP only position, spoiler

aileron 0 Left 0

Rudder 0

IO 8

deg 0 Aileron 1 14 Control

12

surface position, Lt

Lt.

stick position,

Rear deg T: E. UP

Stabilator 0 100 R t.

Rudder force, Ib 0 2

o Stobilator J 12

0 Aileron 1 0 Stabilator Ib

Rt.

Stick force,

Rt Push 12 _I r Transverse o

L LonZdg

UP Pitch 0

-.5 -2 5 0 Roll Angular

Lt. Lt. occet.

Load 5 Normal

I factor

2 UP

0 Pitch

factor 2 load

r Dn

Dn.

Angular 3

velacity, Roll 0 rad /sec

Rt. Rt. I -3 -5 rad/sec2 L YawRtO

Lt. Lt J .5 .8

0 Yaw R t.

-.5 -.8 6

0

6

Nose It. Sideslip angle,

Nose rt. deg IO Nose

UP 0

0 IO 20 30 40 50 60 Time, sec.

Figure 13.- Data obtained during abrupt rolls. Airplane weight, 18,350 pounds; center of gravity at 20.6 percent c; f l i g h t 19.

NACA RM L57E17


Indicated

knots oirspeed,


1.2 I .o .8 Mach .6 number .4 .2

40 spoiler

0

0 Aileron

14 control

Right

UP only position,

Lt. -T stick

Normal

0 IO 20 30 40 50 60 Time, sec.

Figure 14.- Data obtained during full-aileron rolls. Airplane weight, 16,900 pounds; center of grav i ty a t 22.6 percent 5; f l i g h t 19.

NACA RM L5P17 37

Indicated oirspeed, knots

20,000


10.000 600 1.2 500 I .o 400 .8 Mach 300 .6 number 2 0 0 .4 IO0 .2

12

r Rlght

Dn 40 spoiler

0 Left UP only position, oileron 0

surface Rudder O position, Lt 14

IO 8 12

Control 0 Aileron 7 Control stick position,

Lt

TE. Rear de9

UP 0 StabilotorA Stobilotor 0

100 Rt

Rudder force, Ib 0

12

0 Aileron 7 Rt.

Stick

2 0 Stobilotor Ib force,

R t Push r Transverse o 12

Load I

L factor

Angular I L

velocity, rad /sec


! I I I

Lonq. Q I I n. I ,

I .8 - - Rt.

Lt -0 Yaw - - -.8 6

0

6

Nose It.

Nose r t IO Nose

UP 0

0 IO 2 0 30 40 50 60 Time, sec.

_1

Normal load factor

1 1 Angular

accet. rod/sec2

Sideslip angle, de9

Figure 15.- Data obtained during flill-aileron rolls. Airplane weight, 18,600 pounds; center of gravity at 20.3 percent c ; f l i g h t 20.

NACA RM L5Pl7

R e S s U e altitude, f t



1.2 I .o

400 .8 300 .6 200 .4 IO0 .2

12 40 Dn UP

Left aileron 0

Rudder 0 Lt.

0

0

Lt 14

only

Aileron

Mach number

Right spoiler posmon,

7 Control stick IO

8 I2

0 Stobilator_] 12

0 Aileron 1 0 Stobilator force'

12 5

I

2 UP

position,

T. E. UP L Stabilator 0 100 R t. Stick

Rudder farce, Ib 0 2

Rear dyg

Rt.

Rt Push Ibl r Transverse o

L L o n P d g

Load factor 2

UP r Dn. Pitch 0 0 Pitch

Dn. -5 -2

3

velocity, Roll 0 rad /sec

5

0 Roll Angular R t. Rt.

Lt. Lt. -3 -5

L Angle of

attack, deg

.5

Yaw 0

-.5

R t. Lt.

IO Nose

UP 0

Yaw

It.

r t.

2

Normal load factor

1 J Angular

accet. rad/sec*

Sideslip angle, deg

Figure 16. - Data obtained during baxrel rolls. Airplane weight, 17,650 pounds; center of gravi ty a t 22.4 percent E ; f l i gh t 20.

L

I

NACA RM L57E17 39



1 2 ' o o o F F 7,000

400R 300

r Dn

aileron 0 Left

cantral Rudder 0 surface pasltian, Lt

d e l

IO 8

TE.

Stabilator 0 UP

io0 R t.


R t r Trmsverse o Load

factor 2

L L o n q

r UP

Dn. Pitch 0

-.5 Angular

I

velocity, Roll 0 rad /sec -3

Lt. I


I

I

1.2 I .o .8 .6 .4 .2

Mach number

Right spoiler positbn,

i" Control stick position, $4

F i I ' R t .

0 StabilatorA

0 Aileron 1 Stick

Normal load factor

2

0

-2 5

0

-5

UP

Dn.

Rt.

Lt.

Pitch

Roll

' 90 IO0 110 I20

Time, sec

1 Angular

accet. rad/sec2

J Sideslip angle, deg

Figure 16. - Concluded.

40 NACA RM L57E17


lndicoted airspeed, knots

r Control

surface position,

Rudder

r Load

factor

L

Angular I L

velocity, rod /sec

Angle of attock, deg

30,000

25,000

20,000 600 1.2 500 I .o 400 .8 300 .6 200 .4 i o 0 .2

Mach number

Normal load factor

1 J Angular

accel. rad/sec2

Sideslip angle, de9

Figure 17. - Data obtained during rolls. Airplane weight, 17,250 pounds; center of gravity a t 22.2 percent E ; f l i g h t 28.

i

NACA RM L5m17 41





600 500 400 300 number

200 IO0

12

Left aileron 0

Rudder 0

Dn

Lt io 8

TE. UP L Stobilator 0 100

Rudder force, Ib 0 R t.

2

r ~rmsverse 0

R t

Laad factor 2

L L o n T d g

UP Pitch 0

-.5 3

velocity, Roll 0 rad /sec -3 ! yaw Rt.0

.5

r Dn.

Angular R t.

Lt.

Lt. -.5

Nose It. Sideslip


UP 0

Figure 18.- Data obtained during rolling pullout. Airplane weight, 18,400 pounds; center of gravity at 21.6 percent E ; f l i g h t 34.

NACA RM ~ 5 7 ~ 1 7



r Control

surface position,

I8,OOO

13,000 E-

200 ?r IO0 12

Left alieron

Rudder 0

IO Lt

L Stabilator 0

R t. Rudder force, Ib 0 R:r r Transverse o Lodd

factor

Angular velocity, Roll 0 rad /sec -3

R t.

Angle of attack, deg Nos:oL UP 0

0 L

Mach number

2

0"' P i t h

-2 5

0 Roll

-5 .8

0 Yaw

-.8

Dn.

Rt.

Lt.

R t.

Lt

6 Nose It.

Nose rt. r - nn-

I " 0

6 - 1 30 40 50 60

Time, sec

Normal load factor

1 1 Angular

accel. rad/sec2

Sideslip angle, deg

Figure 19.- Data obtained during Immelman. Airplane weight, 16,630 pounds; center of g rav i ty a t 22.4 percent E ; f l i gh t 2.

43



r Control

surface position,

25.000 I I I I 1 I

Rudder force, Ib 0

Normal

factor

Time, sec.

Figure 20.- Data obtained during boost-off f l i gh t , boost turned on a t time 36 seconds. Airplane weight, 17,000 pounds; center of gravity a t 22.0 percent 5; f l i g h t 3 .

44 NACA RM L57E1'7

' Ressure altitude, f t

Indicated airspeed, knots 300

200 IO0

12

I I I

r Left olleron

lot I I I I F - I _1

"" I

Ladd factor 2r*

UP =q L L o n g V g Y "L

Pitch 0 w--

-.5 3

velocity, Roll 0 rod /sec

q

Angular R t

I -3 i

L Angle of

attack, deg

.5

Yaw 0

75

R t.

Lt.

IO Nose

UP 0

0 io 20 30 4 Time, sec.

r 1.2 I .o .8 .6 .4 .2

40

0

0 A i l m

14

UP only

Lt.

I l l 2 R e a r *d I 2Rt.

0 S tab i l a to r l

0 Aileron 7 Stick

Normal load factor

Rt. 0 Roll

-5 Lt.

0 Yaw

-.8

R t.

I 50 60

I Angular occet. rad/sec2

I I

Figure 21.- Data obtained for rudder kick. Airplane weight, 16,400 pounds; center of gravity at 22.0 percent E ; f l i g h t 7.

NACA RM L57E17



r Control


I

lo'oool 5,000 I I 0

1.2 I .o .8 Mach .6 number .4 .2

40 spoiler

0

0 Ailem 7 14 Control

stick 12 position,

Rear deg 0 Stobilator__] 12

0 Aileron 1 OpuShStabilatar pJ 5 Normal

I factor

2 UP

0 Pitch

-2 5 Rt. Roll 1 Angular

Right

UP GnlY position,

Lt.

Rt.

Stick

12 _I

load

Dn.

O Lt. accet. -5 rad/sec2 .8

0 Yaw R t.

Lt 1 6 -.8

Nose It. Sideslip


6

Time, sec.

Figure 22.- Data obtained while opening and closing speed brakes. A plane weight, 16,650 pounds; center of gravity at 21.9 percent E ; f l i g h t 13.

45

.ir -

46 NACA RM ~ 5 7 ~ 1 7



r Control


L Rudder

19,000

600 5 0 0 400 300 200 IO0

Left DnoL olleron

Rudder 0 !s=

Stobilator 0 100

force, Ib 0

r Load

factor

L

Angular I velocity, rod /sec

Rt2r Transverse 0

Pitch 0

-.5 Dn.

Roll 0

-3 L t.

Angle of ottack, deg UP

0-

I I I I "11.2 I .o .8 Mach .6 number .4 .2

40 spoiler

0


stick 12 position,

Rear deg

Right

U P ~ ~ Y position,

Lt.

o Stobilator J 12

0 Aileron 1 OpushStabilator force'

12

Rt.

Stlck

Ibl

5

., I nV

5 0 Roll

-5 .8

0

6 -.8

0

6

Rt.

Lt.

R t.

Lt

Nose

Nose

.8

0

-.8 6

0

6

R t.

Lt

Nose

Nose

Yaw

It.

rt.

20 30 40 50 60 Time, sec.

1

Normal load factor

1 J Angular accel. rad/sec'

Sideslip angle, deg

Figure 23.- Data obtained for loop with roll on top. Airplane weight, l7,800 pounds; center of gravi ty a t 22.3 percent E ; f l i g h t 20.

p

47




1.2 I .o .8 Mach .6 number .4 .2

40 spoiler

0


stick

Rear deg

Right

UP only p w n ,

Control Rudder 0 Lt

12 position,

0 Stobi lotor1 12

0 Aileron 7 OPushStabilotor y’ 5 Normal

I factor

2

0 Pitch UP

-2 5 Rt. Roll 1 Angular

Rt. Stick

12 A

load

Dn.

O Lt. accet. -5 rad/sec2 .8

0 Yaw R t.

-.8 Lt. i 6

0

6


Nose rt. deg

0 10 20 30 40 50 60 Time, sec

Figure 24. - Data obtained during Mach run. Airplane weight, 17,450 pounds; center of gravity at 21.9 percent 5 ; f l i g h t 3 .

48

-~

NACA RM ~ 5 7 ~ 1 7



r Control


2 2 , 0 0 0 d I 1

300k="V: I

I

2 0 0 .4 IO0 I I I I I .2

Mach number

Angular i t

velocity, rad /sec

Rudder

r Load

factor

L

12 40 Right spoiler

Dn UP only position, Left aileron 0

Rudder 0 Lt.

0

0 Aileron 7 Control stick position,

Lt 14 IO 8 12

TE. Rear

UP Stobilator 0

100 R t.

force, Ib 0

deg

o Stobilator J 12

0 Aileron 1 Rt.

Stick

2 OpushStabilator p' Rt 12 _I

Transverse 0

2

5 Normal

I factor

2 . UP

load

LonsfWdg

UP Pitch 0 0 Pitch

Dn. Dn. -5 -2 1

Figure 23. - Airplane

3 Roll 0

5 Rt. I

Rt 0 ~ ~ 1 1 Angular

-3 -5 rad/sec2 L t. Lt. accet.

.5 .8 R t.

Yaw 0 R t.

0 Yaw

-.8 6

Lt. -.5 Lt. I Nose It. Sideslip


IO Nose

UP

6

0

0 IO 20 30 40 50 60 Ttme, sec

f l i gh t 7.

Data obtained during high-speed shallow dive and pullout. weight, 17,400 pounds; center of gravity at 21.7 percent E ;

NACA RM L57E17 49



Control surface position, der

600 500 400 300 number

200 IO0

12 Dn

Left oileron 0

Rudder 0

10 Lt

a TE.

UP L Stabilotor 0 100 R t.

Rudder force, ib 0 2

R t. r ~ronsverse 0 Load

Normal

factor 2

L Long. 2 Fwd

UP

Dn. Pitch 0

-.5 3


.5

Angular Rt.

Lt.

Y0WRt0 Lt. -.5

IO Nose

UP 0

60 70 80 90 100 110 I20 Time, sec

(b)


Angle of ottock, de9

NACA RM ~ 5 7 ~ 1 7


L Angle of

attack, deg

Time, sec.

Nwrnal

factor laad

1 _I Angular

accet. rad/sec2

Sideslip angle, deg

Figure 26.- Data obtained during split-S into Mach run. Airplane weight, 16, goo pounds; center of gravity a t 21.8 percent E ; f l i g h t 14. -

2 0 , 0 0 0 ~ ~

10,000 151000m Ressure altitude, f t


Mach number 300

200 IO0

12 Right

40 spoiler

0

0 Aileron 1 14

UP only pasition, deg

Lt. Control stick

Control surface pos l t i i , deg

Rudder 0 I I

Lt IO 8

TE.

Stabilator 0 UP

100 R t.

force, Ib 0 -

12

0 StabilatorA 12

0 Aileron 1

position,

Rear deg

Rt. Stick force,

OPushStobilotOr Ib I

12

Rudder

r L

Transverse 0 Rt "T 2

Normal load factor

Lodd factor L I Y 2

=oUp Pitch

-2 1 Dn.

Y 5 = Rt. - 0 Roll E Lt. = -5

.e

-

- 0 Yaw -

Rt.

Lt. -.8

1 r Pltch 0

-.5 Dn. I

Angular accet. rad/sec2

I Angular

veloclty, rad /sec

L L t. -3

1 Yaw 0

R t . ' 5 F " t + Lt. -.5

"

6

0 Nose It.

Nose rt.

. "

- 30

Time, sec

Sideslip angle, deg

I P I I I

N o : ? & 4 z 0 60 70 80 !

Angle of attack, deg ZZIkkIT IO0 I10 I20


.. .

NACA RM ~ 5 7 ~ 1 7

airspeed, +vu knots "_F

conwol Rudder o~ ~u 1 surface Control

stick

deg position,

o Stobilator J L Stabilator 0

Rudder force, Ib 0 R t.

.- Rt.

0 Aileron 1 Stick

2 OpushStabilator Ib force,

Rt 12 _I Transverse 0

Load r factor

I

load factor 2 I

2

0

-2 5

Roll 0 0

I _,,F"dn

UP

Dn. -5

3 R t. Rt.

L t. Lt -3 -5

L L", my. r Pitch o Dn.

1 Pitch

Roll

Yaw

Angular

rad /sec velocity,

I


.5 r Rt. F

.8

- 0 r R t.

i Yaw 0 1

j 6 N a s e lt. Sideslip -,5t= I I I I

4 . " 0 Nose rt.

IO z Nose

6

angle, deg

Angle of attack.

0 10 20 30 40 50 60 Time, sec

( a>

.gure 27.- Data obtained during dive t o high speed and high g pull Airplane weight, 16,400 pounds; center of g rav i ty a t 22.0 percent f l i g h t 14.

.out. E ;

-.



r Control

surface position,

Rudder

r Load

factor

L

Angular I I

rad /sec velocity,


i 5 ’ 0 0 0 m

‘oloooE I I 1 I 5,000 I I I I

UP 0

60 70 00 90 100 110 I20 Time, sec

Q ”

53


54

Fi

Flessure altitude, f t


r Control

surface position,

number

Nose rt. deg

.gure 28.- Data obtained during Mach run followed by six consecutive aileron r o l l s . Airplane weight, 17,250 pounds; center of gravity a t 21.7 percent E; f l i g h t 15.

55

Figure



1.2 I .o .8 Mach 6 number .4 .2

40 spoiler

0


12

0 Stobi lator1 12

0 Aileron 1 OpuShStabilator Ib

5 Normol

I factor

2 UP

0 Pitch

-2 5 0 Rt. Roll 1

Indicated

Right

UP only position, de9

Lt

stick position,

Reor deg

Rt.

Rudder force, Ib 0 Stick force,

12 _I

load

Dn.

Lt. accel. -5 radlsec' .8

0 Yaw

-.8 6

0

6

R t.

Lt. 1 Nose It. Sideslip

angli, Nose rt. deg

0 IO 20 30 40 50 60 Time, sec.

29.- Data obtained during Mach run. Airplane weight, 18,750 pounds; center of gravity at 20.1 percent C; f l i g h t 19.




l0,OOo 1 I I

w I I


I I i I I 0 10 2 0 30 40

Time. sec

4 1.2 I .o .a Mach .6 number .4 .2

0 Aileron 1 Control stick

3 I 'Reor d y

position,

7 0 Stobilator1 12

0 Aileron 7 Rt.

Stick

-+)ushStabilator force, Ib I

g>in, Pitch

-2 -jkt, Roll

-5 -1;;; Yaw

3: Nose It.

J

Normal load factor

1 i Angular accel. rad/sec2

Sideslip angle, deg

I 60

Figure 30. - Data obtained during simulated dive bomb run. Airplane weight, 15,450 pounds; center of g rav i ty a t 22.5 percent E ; f l i g h t 14.


1.2 I .o .e Mach .6 number .4 .2

40

0

0 A i l m 1 14 Control

12

0 StabilatwA 12

0 Aileron 1 OpushStabilatcf p' 5 Normal

I factor

UP

Indicated

Right spoiler

UP only position, deg

Lt

stick position,

Rear deg

Rt.

Stick

12 _I

lood

2

0 Pitch

-2 5 0 R t. Roll 1 Angular

Dn.

Lt. accet. -5 rad/sec2 .a 0 Yaw

R t.

-.e Lt. i 6

Nose It. Sideslip


6

Time, sec.

Figure 31. - Data obtained during simulated glide-bomb run. Airplane weight, 19,350 pounds; center of gravity at 19.5 percent E ; f l i g h t

57

16.

NACA RM ~ 5 7 ~ 1 7


r

400 .8 300 .6

.2 200 .4 IO0

12 40

Left aileron O

Dn 0 - . ..

IO0 I I I I I .2 I

Mach number

Right spoiler

surface position.

L Rudder fol

Load r factor

L

Angular i L

veloclty, rod /sec


Ruddel

Stabilata

Stabilatol

rce, Ib

Trmsversa

100 R t.

0 2

R t r O

" Aileron

Stabilata 1

Fwd 2 I

Long.

Pitch 0 0 Pitch

8 2 UP UP

-5 -2

Rt. Roll 0 0 Roll

Lt. Lt. -3 -5

D n Dn.

3 5 R t.

Yaw

It.

rt.

I

0 IO 20 30 40 50 Time, sec

Control I

stick position,

l+J

1 Stick

' Ib force,

J Normal

factor load

1 1 Angular accel. rad/sec2

Sideslip angle, deg

Figure 32.- Data obtained during simulated skip-bomb run. Airplane weight, 18,800 pounds; center of gravity .at 19.6 percent E ; f l i g h t 16.

59



r Control

surface position,

20,000

15.000

l0,OOO 600 1.2 5 0 0 I .o 400 .8 300 .6 200 .4

.2

Mach number

Right 40 spoiler

0

0 Aileron 7 14 Control stick

Rear deg

U P ~ ~ Y positon,

Lt,

12 position,

o Stabilatw J 48

0 Aileron 1 OpushStabllator Ib

5 Notmal

I factor

2

0 Pitch UP

-2 5 0 Rt. Roll 1

Rt

Rudder force, Ib 0 Stick ‘force,

12 _I

load

Dn.

Lt. occet. -5 rad/sec2 .8

0 Yaw R t.

-.8 Lt. 1 6

0

6


Nose rt. deg

0 IO 20 30 40 50 60 Time, sec.

Figure 33. - Data obtained during dive bomb run. Airplane weight, 20,030 pounds; center of gravity a t 19.7 percent E ; f l i g h t 37.

60

R e S S W e altitude, f t


r Control

surface

deg position,


Mach number

.2

40 spoiler

0

0 Ailem 1 14 Control

12

0 Stabi lotwA 48

0 Aileron 1

Right

UP only posmon, de9

Lt.

stick position,

Rear de9

R t

Rudder .force, Ib 0 Stick force,

OpushStabilw Ib

12 _1 5 Normal

I factor

2

0"' Pitch

-2 5 0 Rt. Roll 1 Angular

load

Dn.

Lt accet. -5 rad/sec2 .8

0 Yaw R t.

-.8 u. 1 6

0

6


Nose rt. deg

Figure 34. - Data obtained during dive bomb run. Airplane weight, 18;600 pounds; center of gravity at 21.1 percent E ; f l i g h t 40

NACA m ~57~17 61

Pressure

Indicated

Normal

Time, sec.

Figure 35.- Data obtained during dive bomb run. Airplane weight, 18,300 pounds; center of grav i ty a t 21.8 percent E ; f l i g h t 40.

62 NACA RM L57E17

.

R e S s U e altitude, f t 5,000

0 600 1.2 500 I .o

airspeed, 400 .a Mach knots 300 .6 number

Indicated

200 .4 IO0

12

r surface Lt. 7

.2

40 spoiler

0

0 Aileran

Right

Dn Left UP pasiton, alteron 0

Control Rudder 0

IO position, Lt 14

control stick position,

0 12 TE. Rear

UP Stabilator 0

100 R t.

Rudder farce, Ib 0

r Transverse o

deg

o Stabilatw J 12

0 Aileron 7 Rt.

Stick

2 OpushStaMlatar Rt 12 ‘bl

Load I

L factor

Angular r L

velacity, rad /sec


2

Nwmal load factor

1 i Angular accet. rad/sec2

Sideslip angle, deg

Figure 36. - Data obtained during simulated strafing run. Airplane weight, 19,130 pounds; center of gravity at 19.4 percent E ; f l i g h t 16.


Indicated

knots airspeed,

r I

Control surface

de9 position,

Rudder fol

r Lodd

factor

L

r I

Angular velocity, rod /sec

L


Transverse 0 Normal

Nose It. Sideslip

Time, sec.

Figure 37.- Data obtained during turn into simulated strafing run and pullout into Immelman. Airplane weight, 18,450 pounds; center of gravi ty a t 20.9 percent E ; f l i g h t 18.

64



400 300 2 0 0

I I I I I I

Laad I

factor

L

Angular i L

velocity, rad /sec


2

Lonq?g

Pitch 0

-.5 3

Roll 0

-3

UP

Dn.

R t. L t.

.5

Yaw 0

5

R t. Lt.

IO Nose

UP 0

60 70 00 90 IO0 I

Yaw 0

5 Lt.

Nose UP l o b *

0

60 70 00 90 IO0 I Time, sec

1.2 I .o .8 Mach .6 number .4 .2

I Control

stick 1 I *Rear deg

position,

0 Stabilator I

2 I20

_I

Normal load factor

1 i Angular accet. rad/sec2

Sideslip angle, deg

(b)


"

I



Mach number

l0,OOO

5,000

0 600 1.2

lndicuted 5 0 0 I .o airspeed, 400 .8 knots 300 .6

200 .4 IO0

12

r

L

l- Dn

L Y0WRtO ' Lt. U. 1

.2

40

0 Aileron '1" stick position,

Right

Dn UP only position, alleron 0 Left 0

'On'' Rudder 0

10

spoiler

surface position, Lt 14 Control

deg 8 12

TE. Reor de9

Lt.

UP 0 Stabi latwA Stabilator 0

100 R t.

Rudder force, Ib 0

12

0 Aileron 1 Rt.

Stick

2 OPushStabilator y' r ~r-rse 0

Rt 12 _1

Load 5 Normol

I factor

2

0 Pitch

factor 2 load

L L o n P d g

Pitch 0 UP UP

-5 -2

velocity, Roll 0 0 Lt, Roll A::::,r rad /sec

.5 .8

Dn.

Angular 3 5 R t. Rt. 1 L t. -3 -5 rad/sec2

R t. 0 Yaw

-.5 -.8 6

0

6



UP 0

0 10 20 30 40 50 60 Time, sec.

( 4

Figure 38. - Data obtained during three simulated strafing runs. A i r - plane weight, 17,850 pounds; center of gravity a t 22.2 percent E; f l i g h t 18.

66 NACA RM L37E17


Pressure

60 70 00 90 IO0 110 I20 Time, sec

Figure 38.- Continued.


l0,OOO

5,000

0


200 IO0

aileron 0

Rudder 0 surface position, Lt

I

Rudder

r

TE.

Stabilator 0 UP

100

force, Ib 0 Rt.

2

Transverse 0 Rt

Lodd factor

L

Angular i L

velaclty, rad /sec


IO Nose

UP 0

120 130

Mach number

Right 40

U P ~ ~ Y position, spoiler

0 deg

O Lt. Aileron 1 14

12

0 StobilatorA 12

0 Aileron 1 OpushStobilator F' 12

Cmtrol stick position,

Rear deg

Rt. Stick

I

5

I "V

2

0 Pitch

-2 5

0 Roll

-5 .8

0 Yaw

-.8 6

0

6

UP

Dn.

Rt.

Lt.

R t.

Lt

Nose It.

Nose rt.

3 E E l E L I 0 150 160

Time, sec. 170 180

1

Normal load factor

1 I Angulor

accet. rad/sec2

Sideslip angle, deg


68 NACA RM L5p17

Fi

rad /sec

0 10 20 30 40 50 60 Time, sec

( 4

.gure 39. - Data obtained during over-the-shoulder LABS run. Airplane weight, 19,000 pounds; center of g rav i ty a t 20.3 percent e; f l i g h t 29



r Cantrd

surface position,

l0,OOO I 5,000

0 600 1.2 500 I .o 400 .8 Mach 300 .6 number

2 0 0 .4. ' IO0 .2.

12 40 Right spoiler

Dn Left oileron 0

Rudder 0 Lt. 7 Control stick

UP only posmon, 0

0 A i l m

Lt 14 10 8 12

0 Stabilatoc_]

position,

TE. Rear, deg

L Stabilator 0

Rudder force, Ib 0

UP 48

100 0 Aileron 1 R t

Rt Stick

2 OpushStobilatar p' R t. 12 __I r Transverse 0

2

L L o n P d g

UP [ Dn. -.5 -2

3


j yaw Rt.0

Load 5 Normal

I factor

2 UP

load factor

Pitch 0 0 Pitch Dn.

5 Angular R t. Rt. 1 Roll Angular

Lt. O Lt. -3 -5 rad/sec2

accet.

.5 .8

Lt. u. -.5 -.8

R t. 0 Yaw


6

0

6


IO Nose rt. deg

Nose UP 0

62 72 8 2 92 102 II 2 122 Time, sec


NACA RM ~ 7 7 ~ 1 7



r Control


l0,OOO I 1 I c

t " - - -LHp I 400 300 200 A IO0

Mach number

Right 12 Dn.

oileron 0 Left

Rudder 0 L t

10 8

TE. UP L Stabilator 0 100 Rt


r Transverse o Rt

I Load

foctor 2

L LongfWdg

UP [ Dn. Pitch 0

-.5

Angular R t. 3

velocity, Roll 0 Lt. rod /sec -3

.5

YowRtO Lt. -.5

IO Angle of Nose

UP attack, deg 0

0 I O 20 30 40 50 60 Time, sec

Figure 40.- Data obtained during over-the-shoulder LABS run.

Normal load foctor

1 I

Angular occet rod/sec2

1 Sideslip angle, deg

Airplane -weight, 20,030 pounds; center of gravity a t 19.7 percent E ; f l i g h t 29.



Rudder 0

Nose It. Sideslip

6 2 7 2 8 2 92 102 112 122 Time, sec

(b)



R e S s U e altitude, f t


r Control

surface posltkm. d y

I .2 I .o .8 .6 .4 .2

40

0

0

14

UP

Lt.

only

Aileron

Mach number

Right spoiler

’i” - 9

control stick

r Load

factor

IO 8 12 position,

TE. Rear deg o Stbbilator J

Stobilator 0

IRudder force, Ib 0

’ UP 48

L 100 0 Ailyon zck

2 opushsta~~atw ;FJ Rt

R t.

R t Transverse 0

12

I “T I I I A5 2

UP .I

Pitch 0

-5 3

vdoclty, Roll . .O Angulqr Rt.

Roll rad /KC

I Angle of

attack, deg

.5 .8

Lt. Lt.

Rt. R t. Yaw 0 0 Yaw

-.5 6 -.8

Nose It.

0 I

0 IO 20 30 40 50 Time, sec.

60

A Narrnal

load factor


Sideslip angle, de9

Figure 41.- Data obtained during simulated over-the-shoulder LABS run carrying 1,700-pound practice bomb. Airplane weight, 18,6510 pounds; center of gravi ty a t 20 .O percent E ; f l i g h t 48.

73



r Control


- - -

Vi -

HP 0

600 - 500 / 400 -- 300 -

1.2 -

.e - 1.0

.2

.4 -

.6 -

-

- -

- M -

200 1 - IO0

12 40 Dn Up only Left 0

aileron O

Rudder 0

0 Aileron Lt.

Lt bomb r e l e a s e 14

Mach number

Right spoiler position,

i’ Control

stick

Nose It. Sideslip

74 NACA m ~ 5 7 ~ 1 7



r Contra1


12 40 Dn

Left aileron O

Rudder 0 Lt.

UP 0

0

14 Lt

12 40 Dn UP

Left aileron O 0

0 Lt.

only

Aileron

Mach number

Right spoiler

'i" m n ,

Control stick

10 8 12 position,

Rear deg TE. UP

Stabilator 0 100 Rt


o Stabilator J 48

0 Aileron 1 0 ,,shstabilatar

R t

Stick

I r ~ronswrse 0

Rt 12

5

I b o d

factor 2

L Long"! 2

UP [ Dn. Pitch 0

-.5 3


UP 0 Pitch

-2 5

0 Roll

Dn.

Angular R t. Rt.

Lt. Lt. -7 -5 L yaw Rt':

.5

Lt. 75

IO Angle of Nose

UP attack, deg 0

1 I I I I I I 0

Time, sec. IO 20 30 40 50 60

2

Normal load factor


Sideslip angle, deg

Figure 43.- Data obtained during over-the-shoulder LABS run dropping a l,7OO-pound practice bomb. Airplane weight, before dropping bomb, 19,530 pounds; center of gravity at 19.9 percent with bomb, 20.2 percent E a f t e r dropping bomb; f l i g h t 53.

75



r Control


Angle of ottack, de9

l0,OOO I

600 500 400 300 2 0 0 I O 0

1.2 I .o .8 .6 .4 .2

Mach number

Right 40 spoiler

0


12

0 StabilatorJ 40

0 Aileron -1 0 Stabilator

UP only pasiton,

Lt.

stick position,

Rear deg

Rt

Rudder force, Ib 0 Stick

Push 12 _I 5 Nwmal

I factor

2

0 Pitch UP

-2 5 0 Rt. Roll 1 Angular

Lt. accet.

load

Dn.

-5 rad/sec2 .8

0 Yaw

6 -.8

R t.

u. 1 Nose It. Sideslip


6

60 70 80 90 IO0 110 I20 Time, sec

(b)


NACA RM L57E17



r Control

surface position,

L

r Rudder

10,000

5,000

0 600 5 0 0 400 300 200 IO0

Mach number

Lodd factor

L

Angular r rad /sec velocity,

L Angle of

attack, de9

. "

12 Right

Dn spoiler

Left aileron 0 deg

Rudder 0

UP only pasiton,

Ailemn 1 Lt Control

stick position, IO

8 TE.

UP Stobilator 0

100 R t.

force, Ib 0 2

Tronsverse 0

deg

Stobilator J Aileron 1 Stablotor force'

Stick

Rt 1 9

2

Longtwdg

Pitch 0 Pitch UP Dn. -5

3 R t.

Roll 0 Lt. -3

Roll

2

Longtwdg

Pitch 0 Pitch UP Dn. -5

3 R t.

Roll 0 Lt. -3

Roll

.5

Yaw 0

-.5

Rt.

Lt.

IO Nose

UP -

=/- 0 "

0 I O 20 30 40 50 60 I

Time. sec

"I

Normal load foctw

1 J Angular accet. rad/sec2

Sideslip angle, deg

Figure 44.- Data obtained during initial-point LABS run. Airplane weight, 18,500 pounds; center of gravity a t 21.3 percent E ; f l i g h t 31.

77

Ressure

I .2 I .o .8 Mach .6 number .4 .2

40 spoiler

0


stick

Rear deg

Right

UP only posimn,

Lt.

12 position,

o StabilatorJ 48

0 Aileron gck 0 Stabilator pt

Rt

Push 12 _I

Laab factor

L

Angular 1 L

velaclty, rod /sec


Normal load factor

1 1 Angular

accet. rad/sec2

Sideslip angle, deg




r control

surface

deg position.

I

I HP 0

600 1.2 500 \ 1.0 400 -" Mach 300

- - Vi - - -

- M - - rumber

200 1 IO0 .2

- - .4 - -

12 Dn

40

0

0 Aileron

Left aileron 0

Rudder 0 Lt.

UP only

Lt 14

Right

i" spoiler position,

control stick

10 8

TE.

R t. Rudder force, ib 0

2 R t r Tronsverse o

Load 2 factor L Pitch 0

-.5 3


UP

Angular R t. Lt.

Lt. -.5

Nose It. Sideslip

IO Angle of Nose

UP attack, de9 0

0 IO 20 30 40 50 60 Time, sec.

Figure 45.- Data obtained during simulated initial-point LABS run carrying 1,700-pound practice bomb. Airplane weight, 20,430 pounds; center of gravity a t 18.2 percent c ; f l i g h t 30.



r Control

surface position,

Mach number

12 Dn.

40

Left Up only

aileron 0 0

0 Ailerm Rudder 0:

- . .

8, _c

Lt. Lt 1 14 -

IO - Bom,b sw i t ch on+

Right

7 spoiler p W n ,

Control stick position, d?

L UP

100

Rudder force, Ib 0 R t. Stick

0 Stabilatw_]

0 Aileron 1 Stabilator 0 48

Rt

2 OpushStabilator p' Rt r Transverse o 12

A

Load factor

i

Angular 1 !

velocity, rad /sec

Angle of ottack, deg

Figure 46. -

5 Normal

factor 2 load

Long?j I

Pitch 0 0 Pitch Dn. Dn.

UP 2

UP

-.5 -2 1 R t. 3 5

Rt. I Roll 0 0 Roll Angular

Lt. Lt. accet. -3 -5 rad/sec2

.5 .8 R t. R t

Lt. u. -.5 -.8

Yaw 0 0 Yaw

angle,

I 16NOSe rt. deg

a

0 10 20 30 40 50 60 Time, sec.

Data obtained during initial-point LABS run dropping a

79

Indicated

knots oirspeed,

number

I

L Load

factor

r L

Angular

rad /sec velocity,


Pitch

Roll

.5

Yaw 0

-.5

Rt.

Lt.

.8

0 Yaw

-.8

R t

Lt

I I I I I I

Nose &\ "16 UP

0 I 1 I

6

0

6

Nose It.

Nose r t IO Nose

UP 0

60 70 00 90 IO0 I10 I20 Time. sec.

I I I I I I I 60 70 00 90 IO0 I10 I20

Time. sec.

Normal load factor

1 1 Angular

accet. rad/sec2

Sideslip angle, deg

(b)


Fi

Ressure oltitude, f t


r Control


- - I -

15,000 I 600 - 5 0 0

- - 1.2

-

400 1.0

.4 "

.6 - - vi - 300

.8 - - -

- - M

-

2 0 0 = - -

IO0 -

.2

- - I -

15,000 I 600 - 5 0 0

- - 1.2

-

400 1.0

.4 "

.6 - - vi - 300

.8 - - -

- - M

-

2 0 0 = - -

IO0 -

.2

40 Up only

0

0 Aileron

14 Lt.

Mach number

Right spoiler posihn,

7 Control stick

I I I 1 I 1 Normal

Loa'd factor

i

r Angular

rad /sec velocity,

Angle o f attack, deg

2

LonsfWdg

UP Pitch 0

-. 5 Dn.

load factor

Rt. -

-3

3: 5 Roll 0 -,-. .-

Rt.

Lt. 0 Roll -

Lt. -5

Rt.

Lt.

.5 .8

Yaw 0 R t.

Lt 0 Yaw

-.5 -.E 6

0

Nose 6 IO Nose rt.

UP. 0


I I I I I I I 0 IO 2 0 30 40 5 0

Time, sec 60

angle, deg

.gure 47.- Data obtained while ra t rac ing ( s ta l l - tu rns and pitch-up) Airplane weight, 13,930 pounds; center of g rav i ty a t 22.1 percent f l i g h t 14.

81

82 NACA FM L57E17



r Contrd


1 Rudder

r

I5.000 600

Mach 300 number 200

I I I i i 1.2

Lodd factor

L

Angular r i

velocity, rad /sec


Right 40 spoiler

0

0 Ailwon 7 14 Control

stick

Rear deg

U P ~ ~ Y positon,

Lt.

12 position,

o StobilatorJ 12

0 Aileron 1 o Stablatar :?

Rt.

Stick

Push Trmsverse 0 12

60 70 00 90 IO0 I IO I20 Time, sec

(b )

-I

Normal load factor


Sideslip angle, deg

Figuse 47.- C o n c l u d e d .


30,000

I I i 1


r Co4rol


L

r Rudder

6mk 500 I 400

- 300

- - M

- .6 - .8 - -

-

200 .4 - vi-" - 7 -

1001 I I I I I I .2

0

14 Rudder 0

Lt Lt.

IO

only .

Aileron

Mach number


Control stick position, dqg

. - 8 12

TE. Rear

UP 0 StobilatorA Stobilotor 0

100 R t.

force, Ib 0

12

0 Aileron 1 Rt.

Stick

2 OpuShStobilator rJ Rt

Transverse 0 12

I Lodd

factor I L

r I

Angular velocity, rad /sec

1 Angle of

attack, deg

_I

Normal load factor


Sideslip angle, de9

Figure 48.- Data obtained while rat racing. Airplane weight, 18,430 pounds; center of g rav i ty a t 20.0 percent E ; f l i g h t 16.

84



r Control

surface position,

25,000

20,000

15,000 600 1.2 500 I .o 400 .8 Mach 300 .6 number

200 .4 IO0 .2

Right 40 spoiler

0


stick

12 position,

Rear deg

UP only position,

Lt.

o Stabilotw J 12

0 Aileron 7 Stabilator Ib

'Push A 12 5 Normal

I factor

2 UP

0 Pitch

-2 5 0 ~ ~ 1 1 Angular

Lt. accel.

Rt.

Stick force,

load

Dn.

Rt.

-5 rod/sec2 .8

0 Yaw R t.

Lt

6 -.8

Nose It. Sideslip


6

0 IO 20 30 40 50 60 Time, sec.

Figure 49.- Data obtained while rat racing (flew through j e t wash a t time 2 seconds) with pitch-up. Airplane weight, 17,300 pounds; center of g rav i ty a t 22.0 percent E ; f l i g h t 16.



25,000

20,000

15,000 600 1.2 500 I .o 400 .8 300 .6 200 .4 IO0 .2

Mach number

Loid factor

L

Angular r L

velocity, rad /sec


.8

0

-.8 6

0

6

R t.

Lt

Nose

Nose

- -

60 70 80 90 IO0 I20

Yaw

It.

r t.

J Sideslip angle, deg

Time, sec

(b)


I I

NACA RM L5’j%1’7



r Control

surface position,

Mach number

120 130 140 150 160 170 180 Time, sec.

( e >


NACA RM L5P17



r Codtrol


20,000

15.000

10,000 600 500 400 300

1.2 I .o .8 .6

200 .4 I O 0 .2

Mach number

Right spoiler

control stick

dqg position, 8 12

T E. 0 StabilatorA UP 12

100 Rt. R t. 0 Aileron 7

Rear

Stobilator 0

Rudder force, Ib 0 Stick 2 OpushStaMlator

Rt r Transverse o 12 "1

Load I

factor I L

r I


I

L Angle of

attack, deg

Normal

factor load

1 Angular

accel. rad/sec2

Sideslip angle, deg

Figure 50.- Data obtained while rat racing. Airplane weight, 17,070 pounds; center of gravity at 22.1 percent E ; f l i g h t 16.

88



r Control

surface position,

L r

Rudder

20'ooog-1

12 40 Dn.

Left oileron O

Rudder 0 Lt.

UP 0

0

Lt 14

only

Aileron

Mach wmbw


'i" control stick

100 I R t. Stick

force, Ib 0 2

Transverse 0

force, 0 PushStobilator Ib

' R t 12 _I

Lodd factor

L

Angular i i

velocity, rod /sec


lood factor

2 UP UP

Pitch 0 0 Pitch Dn. Dn. -.5

3 Roll 0

-2 5 Rt. R t. 0 Roll Angular

occet. Lt. Lt. -3 -5 rod/sec2

.5 .a R t. Lt.

R t.

u. Yow 0 0 Yaw

-.5 6 -.8

0

6

Nose It.

Nose rt. IO

Nose UP

0

6

I Sideslip angle, deg

Time, sec

(b)


NACA RM L57l317

I


I I


Load factor 2

L L o n P d g

Pitch 0 UP

-.5 r Dn.

I Angular

velocity, rad /sec

t Angle of

attack, deg

Nose UP l o r 0

120 130 140 150 Time, sec.

T I .2 I .O .8 .6 .4 .2

Mach number

Right spoiler position, deg -

12 .

0 S tab i l a to r l 12

0 Aileron

OPushStabilatar force'

12

Rear

Rt. Stick

'b,

I Control stick position, d;g

2

0 Pitch UP

-2. Dn.

5 Rt.

O Lt. Roll

-5 .8

0 Yaw

-.8

R t.

Lt.

6

0

6

Nose It.

Nose rt.

0 170 I80


J

Normal load factor

1 Angular accet. rad/sec2

Sideslip angle, deg

I I I I I

NACA RM L57E17

Time, sec

Figure 51.- Data obtained during pitch-up followed by Mach run. Airplane weight, 17,150 pounds; center of gravity at 22.2 percent E ; f l i g h t 2.



r Cantrd

surface pasitlan,

300 200 I00

12 Dtl

aileron 0 Left

Rudder 0

io Lt

8 TE. UP L Stabilator 0 io0 R t.

Rudder force, Ib 0 2 r Transverse o

R t

Load factor L

Angular I L

velocity, rad /sec

2

Long. 9 Fwd

UP

Dn Pitch 0

-5

Roll 0 -3

Lt.

Angle of

deg attack,

I I I

0 IO 20

Mach Rmber

Roll

-5

40 50 60

Figure 52.- Data obtained during pitch-up. Airplane weight, 16,900 pounds; center of gravity at 21.8 percent E ; flight 5.

L Angle of

attack, deg

.5 = .8 Rt. 1

Lt.

Rt.

Lt -.8

Yaw 0 0 Yaw

-s - 6

0

6

Nose It.

Nose rt.

\

I 0 I O 20 30 40 50 60

Time, sec

J Normal

load factor

1 1 Angular

accet. rad/sec2

Sideslip angle, deg

Figure 53.- Data obtained during pitch-up. Airplane weight, 17,550 pounds; center of grav i ty a t 21.5 percent E ; f l i g h t 9.

93

' Ressure altitude, f t 15.000

10,000


r Control

surface position,

number

L Angle of

attack, deg

Rt. Yaw 0 0 Yc

Lt. Lt -.5

.3 .V

R t.

-0

6

a 10

Nose UP

6

0

0 10 20 30 40 50 60 Time, sec.

I I I I I

IW

.V

Nose It.

Nose rt.

Normal load factor

1 1 Angular

accet. rad/sec2

Sideslip angle, deg

Figure 54.- Data obtained during Immelmaa followed by pitch-up. Airplane weight, 17,400 pounds; center of gravity at 21.7 percent E ; f l i g h t 11.

6;

94

Pressure altitude, ft

NACA RM L5Pl'j'

32,000

27,000

22,000 600 1.2

Indicated 500 I .o

airspeed, 400 knots 300 .6

.8 Mach number

200 .4 IO0 .2

r 12 Right

40 spoiler

0


stick 12 position,

Rear deg

Left aileron 0

Rudder 0

IO 8

Dn. U P ~ ~ Y poston,

surface Lt.

position, Lt

'1 TE.

UP o Stobilator J

Stabilator 0 12

100 0 Aileron 1

2 OpushStabilotor 12 1

Rt.

Rudder force, Ib 0 R t. Stick

R t r ~ r m s e r s e 0 Load

5 Normal

I factor

2 UP

factor 2 load

L Long?dg

UP Dn.

Pitch 0 0 Pitch Dn. -5 -2

3


L yaw Rt.0 Lt. Lt 1

5 Angular R t. Rt. 1 Roll Angular

L t. O Lt. -3 -5 rad/sec2

accet.

.5 .8 R t.

0 Yaw

6 -.8 -.5

Nose It. Sideslip


IO Angle of Nose

UP

6 attack, deg 0

0 IO 2 0 30 40 50 60 Time, sec.

( 4

Figure 55. - Data obtained .during stall, a s ta l l turn, and then pitch- Airplane weight, 17,000 pounds; center of grav i ty a t 21.8 percent f l i g h t 15.



32,000

27.000

22,000 600 1.2 500 I .o 400 .8 300 .6 200 .4 IO0 .2

I

Con‘trol surface position.

Mach number


factor

60 70 80 90 100 110 I20 Time, sec

(b)


95



Control r de'

surface position,

24'0mm 19,000

1.2 I .o

400 .8 Mach 300 .6 number

200 .4 IO0 .2

12 Right

Dn 40 spoiler

0

0 Aileron '1" stick position,

UP only posmon, Left aileron 0

Rudder 0 Lt, Lt 14 Control

8 12 TE. Rear

IO

deg

UP 0 StabilatorA

2 OpushStabilator force' Rt 12 Ibl

L Stabilator 0 100 R t.

Rudder force, Ib 0

12

0 Aileron 1 Rt.

Stick

r ~ r m s e r s e 0 Load

5

I

2

0 Pitch

factor 2

L Longfwdg

Pitch 0 UP r Dn.

UP

Dn. -5 -2

3 5 Rt. Rt. Lt. Lt.

Angular velacity, Roll 0 0 Roll rad /sec

I -3 -5

! Angle of

attack, de9

.5

= " Yaw 0

r Rt. .e Rt. 1 / i E - Lt. =

1- - - r 8 <- - -0 Yaw

-.8 - -.5 Lt.

6

0

6

Nose It.

Nose rt. IO Nose

UP

0 10 20 30 40 50 60 Time, sec.

2

Normal load factor


Sideslip angle, de9

Figure 36.- Data obtained during pitch-up. Airplane weight, 17,100 pounds; center of g rav i ty a t 22.3 percent E ; f l i g h t 19.

NACA RM L57E17

25,000

Ressure altitude, f t 20,000


400 300 200 IO0

aileron 0

Rudder 0 surface position, L t deq

Rudder

r force, Ib 0

Rt Transverse 0

Looh factor L L o n p d 8

Pitch 0 r UP Dn. -.5 2 F -

Angular I

velocity, rad /sec

t Angle of

attack, deg

Roll 0

-3 L t.

I O L -

0 I

I

Right

- deg wj:Lt Aileron I Control stick

12 position, Rear deg

0 Stobilator_] 4 8

0 Aileron 1 0 Stabilator force'

Push Ibl 12

Rt Stick

Dn. -2

Rt.

Lt. I 1 4 - 5

Pitch

Roll

-.8

Nose It.

Nose r t

6 7

97

Normal load factor

1 I

Angular occel rod/sec2

Sideslip angle, deg

Figure 57.- Data obtained during pitch-up. Airplane weight, 17,300 pounds; center of grav i ty a t 22.2 percent E ; f l i g h t 51.

NACA RM L57E17

F'ressure altitude, f t


r Contrd

surface position,

300 200 IO0

1 2 t I I

TE.

UP L Stabilotar 0 100

Rudder farce, Ib 0 R t.

Load I

factor

L

Angular r L

velocity, rad /sec -3

R t.

Lt. Yaw 0 ._


$ .4

Mach number

0 Aileron

14

Y I 12

Lt. 1- Control stick position,

Rear de9 - 0 StabilatorA

7 0 Aileron 1 -= 48

Rt

Stick mr Stabifator

force, Ibl

2

Normal load factor

1 1 Angular. accel. rad/sec2

Sideslip angle, de9

Time, sec


Figure

Rear

10

T.E. up

8

6

4

2

0

2

4000

2000

0

2 -2OOG

2 $ -4OOL

5 Down

- dOOC

1

4000

e_ 0 c .+. 1

% 2 -4OCC E a, tn

3 -8000

2 -12000

9 Y T.E. up

-16000 0 1 2 3 4 5 8 I 8 9

Time, seconds

99

58.- Calculated horizontal t a i l loads and hinge moments obtained during the pitch-up of figure 57.

I

100 NACA RM ~ 5 7 ~ 1 7

Ressure oltitude, f t


. r Control

surface position,

L

r Rudder

Load factor

L

Angular 1 I

velocity, rad /sec


15,000

10,000

5,000 600 1.2 500 I .o 400 .8 Mach 300 .6 number 200 .4 IO0 .2

12 Right

Dn 40 spoiler

0

0 Aileron 7 stick position,

Left UP only position, aileron 0

Rudder 0 Lt. Lt 14 Control

IO 8 12

TE. Rear deg

UP 0 StabilotorA Stabilotor 0

100 Rt

12

0 Aileron 1 Rt.

force, Ib O ~ w ~ w , ~ ~

Stick

2 t Fs OpushStabilator p' Transverse 0 -

R t "T 12 A I I I I I I I

2

L o n g Y g I

Pitch 0 UP

2 UP

Dn. Dn. -.5 -2

Pitch

3

Roll 0 R t. L t. Lt. -3 -5

5

0 Roll Rt.

.5 R t.

Lt.

= .8 Yaw 0 "

i Rt.

: u. - - ~~11 r 0 Yaw

-.5 - -.8 6

0

6

Nose

Nose IO Nose

UP 0

0 IO 20 30 40 50 60 Time, sec.

It.

r t.

Normal load factor

1 J Angular accet. rad/sec2

Sideslip angle, deg

Figure 39. - Data obtained during ILD (instrument let down) and GCA (ground control approach) i n formation. Airplane weight, 18,950 pounds; center of gravi ty a t 19.5 percent E ; f l i gh t 17. -


101

Indicated

knots airspeed,

c

0 I I I I

400 Mach 300 number 200 M IO0 .2

12 Right

I Dn. 40 spoiler

0


Left UP only position, aileron 0

Rudder 0

IO

surface Lt.

position, Lt 14

8 12 T E. Reor

deg

UP 0 StabilatorA Stabilator 0

100 R t.

Rudder force, Ib 0

12

0 Aileron 1 Rt.

Stick

2 0 Stabiiator pJ Rt Push

r Transverse o 12

I Load

factor

L

Angular r velocity, rad /sec


I I I I I I I

Rt. R t.

Lt. Lt. 75 -.8

.5 .8

Yaw 0 0 Yaw

6

0

6

Nose It.

Nose rt. IO Nose

UP 0

60 70 80 90 IO0 110 I20 Time, sec

( b )


1 Normal

load factor

I Angular

occet. rad/sec2

1 Sideslip angle, de9

102



r Control


L Rudder

r

500 400 300

Mach number

200 M

I I I I I .2

I

L Load

factor

Angular I rod /sec velocity,


12 Right Dn

40

0


Left oileron 0

Rudder 0

spoiler UP only position,

Lt Lt.

IO stick 8 12 position,

TE. Rear deg

UP 0 StabilatorA

2 OpuShStabilotor force' Rt 9 2 Ibl

Stobilator 0 12 100 R t.

force, Ib 0

Rt. 0 Aileron 1

Stick

Transverse OF -

1

I

5 2

-."Sfwdg I

UP 2

0 Ditch 0 Dn,

UP

Dn. - -3 I 1 -L

R t. 3: .,I, I P ._ I ,~.. Roll 0

Lt. r -3

5

0

-5

Rt.

' I p I . -

Lt. -

Pitch

Roll

.5 .8 R t. R t

Lt. U. -.5 -.8

Yaw 0 0 Yaw

6

0

Nose 6

Nose It.

Nose rt. IO

UP " -

0

I20 130 140 150 160 170 180 Time, sec.


2

Normal load factor

1 J Angular accet. rod/sec2

Sideslip - angle, de9

NACA RM L57E17



(dl


104 NACA RM ~ 5 7 ~ 1 7



r Control

surface position,


I I 1.2 Broke open

400 .8 300 - . v i

-

.2 .

.4 -

.6 206 1

-

IO0 M

Normal

240 250 260 270 280 290 300 Time, sec.

( e >

Figure 39. - Continued.

Mach number

"

NACA RM L57E17


Indicated

knots airspeed,

- - j 0

6001

- - 300 .8 - - 400 1.0 - ""

500- 1.2 - - Broke 'open

.6

.4 IO0 M .2

-1 - - - -

200 Vi 1 - -

Mach number

control surface position, de9

I

L

r

Rudder

I

L Load

factor

r Angular

rad /sec velocity,

I

12 Right 40

0

0 Aileron 7 stick position,

Dn. spoiler Left UP only position, olleron 0

Rudder 0 Lt. Lt 14 Control

IO 8 12

T E. Rear

UP

deg o Stabilotor J 12

0 Aileron 1 Stabilotor 0

100 R t.

force, Ib 0

Rt.

Stick force,

2 0 Stobilator Ib Rt Push

Transverse 0 12

I I I I I I 2

I

2 UP

0 UP

Dn. Pitch 0

Dn. -.5 -2 3 5

R t. L t.

Roll 0 Rt.

0 Lt.

-3 -5

UP

Dn. Pitch 0

UP 0

Dn. -.5 -2 3 5

R t. L t.

Roll 0 -3

Rt. 0

-5 Lt.

Pitch

Roll

.5 L Yaw 0

.8

0 Yaw Rt. R t.

Lt. Lt. -.5 -.8

6

0

Angle of Nose 6

Nose It.

Nose rt. IO

attack, deg UP

0

300 310 320 Time, sec.

330 340 350 360

Normal

factor load

1 I

Angular occet. rad/sec2

Sideslip angle, des

(f>


106



r Control

surface position,

1.2 I .o .8

300 .6 200 .4 IO0 .2

Mach number

." 12 Right

Dn. 40 spoiler

Left UP only position, olleron 0 0

0 Aileron 7 Rudder 0 Lt.

stick position,

Lt 14 Control

TE. Rear

IO 8 12 deg

UP 0 StabilatorJ L Stabilator 0 100 R t.

Rudder force, Ib 0

r Transverse o

L LonsfWdg

Pitch 0

12

0 Aileron 1 Rt.

Stick

2 0 Stabilator p' Rt Push

12 5

I

2

0 Pitch

Load factor 2

r UP UP

Dn. Dn. -5 -2

I Angular 3 z 5

Rt. - P Rt. velocity, Roll 0 - - - -. rad /sec Lt. - P

I

O Lt. Roll

-3 - -5

L Angle of

attack, deg

- Rt. Y" "" ." - \I I - - 0

5 Lt. - -.8 6

0

6

Nose

Nose 10 Nose

UP 0

I I I I I I I 360 370 300 390 400 410 420

Time, sec.

Yaw

It.

r t.

J Normal

load factor

1 J Angular accet. rod/sec2

Sideslip angle, deg

( g )



Indicated

knots airspeed,

r Control

surface position,

L Rudder

r

5,000 - I

0 600 - - 1.2 500 - - 1.0 400 - -- .8 300 - .6 -

2 0 0 = - .4 IO0 .2

Mach number

Lodd factor

r L

Angular

rad /sec velocity,

L Angle of

attack, de9

Right 40 spoiler

0


UP only position, de9

Lt.

stick

Rear deg 12 position,

o Stabilator J 12

0 Aileron 1 OpushStabilator force'

Rt.

force, Ib 0 Stick

Ib, Transverse 0 12

5 2

Fwd._ I 2

0 Pitch UP

3: I Rt -

Roll 0 Lt. -3

.5 =

Y 5 * '0 Roll

E Lt. -5

, P . , I Rt.

- =

UP 0

A

Normal load factor

1 I


Sideslip angle, deg

I I I I I I 0 IO 2 0 30 40 50

I Time, sec.

60

Figure 60.- Data obtained during formation flying. Airplane '

17,430 pounds; center of gravi ty a t 22.3 percent E ; f l i gh t 25. weight ,

108 NACA RM L5P17

Fressure altitude, f t


r Control

surface position,

L Rudder

r Ladd

factor

L

Angular r i

velocity, rad /sec


Mach number

200 I on ."

12 40 Right

Dn UP only position, spoiler

aileron O Left 0

Rudder 0 Lt. 0 Aileron 7

Control stick position,

Lt 14 IO 8 12

TE. Rear deg .

UP 0 Stabi iotwA Stabilator 0

100 R t.

farce, Ib 0

12

0 Aileron 1 Rt.

Stick

2 OpushStabilatw r1 Rt 12 A

Transverse 0 5 Normal

2

Long?$

Pitch 0 UP

-.5 -2

I

2

0 Pitch UP

Dn. Dn.

3 Roll 0

5

0 Roll Rt.

L t. Lt. -3 -5 .5 .8

Rt. R t.

Lt. Lt 75 -.8

Rt.

Yaw 0 0 Yaw

6

0

6

Nose It.

Nose rt. IO Nose

UP 0

60 70 80 90 IO0 110 I20 Time, sec

factor load

I 1 Angular

accet. rad/sec2

Sideslip angle, deg


1.2 I .O .8 Mach .6 number .4 .2

40 spoiler

0


Indicated

Right

UP only position, deg

Lt.

stick

Rear deg 12 position,

0 StabilatorA 48

0 Aileron 1 0 Stobilator

12

Rt Stick

Push I bl

Load I

factor

1 L

L



5 2

L o n P d g I

UP

Dn. Dn. -.5 -2

2

0 Pitch Pitch 0 UP

3

Roll 0 R t.

5

0 Roll Rt.

Li. Lt. -3 -5

Rt. .5 .8

R t.

Lt. u. Yaw 0

-.5

0 Yaw

6 -.8

0

6

Nose It.

Nose rt. IO Nose

UP 0

0 IO 2 0 30 40 50 60 Time, sec.

( a) , .

J

Normal load factor

I 1 Angular accet. rad/sec2

Sideslip angle, deg

Figure 61.- Data obtained during in-flight refueling. Airplane weight, 19,900 pounds t o 25,450 pounds; center of grav i ty a t 20.0 percent S. at contact, ranging from 20.0 t o 17.7 t o 20.6 during refueling; f l i g h t 52.

110 NACA RM L57E17



r Control

surface position,

15,000

10,000

5,000 600 1.2 500 I .o 400 .8 300 .6 200 .4 IO0 .2

L 100 R t.

Rudder force, Ib 0 2 r Transverse 0

R t


Mach number

Normal

60 70 00 90 IO0 110 I20 Time, sec

( b )




r I

Control surface position, de,g

L Rudder

r Load

factor

r L

- -

5,000

600

300 -

- 400

- 5 0 0

- -

-

- 1.0 - 1.2

.8 - .6

-~~~ .4 .2

- - -

-

2 0 0 = Vi - - 100 - M

5 ~ z ~ . ~

1 .O 400 5 0 0 .8

12 Dn

40 Left Up only aileron 0 0

Rudder 0 0 Aileron

Lt 14 Lt.

Mach number

Right spoiler position, deg

Control stick position,

Angular I

velocity, rad /sec

"(5


Long.

UP Pitch 0 0 Pitch

Dn. Dn.

UP

-2

"

Pitch 0 2 . -. . - - "-4 - -. -. UP g 4 Dn. =

- - 0 Pitch

-5 = - -2 : Dn.

3 Roll 0

R t. Lt.

5

0 Roll Rt.

Lt. -3 -5

.5 .8 Rt.

Lt. Yaw 0 0 Yaw

R t.

u. -.5 -.8

6

0

Nose 6

Nose It.

Nose rt. 10

u p o r = 120 130 140 150 160 170 180

Time, sec.

Normal load factor

1 I

Angular

rad/sec2 accet.

Sideslip angle, deg

( 4


112

I

NACA RM L57El'j'

Fressure altitude, f t


r ""

Control surface position,

15,000

10,000

5,000 600 1.2 500 I .o 400 .8 Mach 300 .6 2 0 0 .4 IO0 .2

number

."

12 Right 40

0


Rear deg

Dn spoiler

Left aileron 0

Rudder 0 Lt.

UP only position,

L t 14 10 8 12

TE. UP

o Stabilator J L Stabilator 0

Rudder force, Ib 0

48 100 0 Aileron 1 Rt R t. Stick

2 OPushStabilatcf ET1 Rt r Tronswrse o 12

5

I

2

0 Pitch

Load factor 2

L Long"$

Pitch 0 UP

-.5 -2 r Dn.

UP

Dn.

Angular 3 5 R t. Rt.

Lt. Lt. -3 -5

Velocity, Roll 0 0 Roll rad /sec

L Angle of

attack, deg

.5 .8 Rt. R t.

Lt. Lt. -.5 -.8

Yaw 0 0 Yaw

6

0

6

Nose It.

Nose rt. IO

Nose UP

0

180 190 200 210 220 Time, sec.

230 240

_I

Normal load factor

1 1 Angular accet rad/sec2

Sideslip angle, deg

(dl

Figure 61. - Continued .

NACA RM L57E17

15,000



2 0 0 IO0

I Control


Rudder 0 L t

10 8

TE.

Stabilator 0 UP

100

Rudder force, Ib 0 Rt

2

O R t. r

Load factor 2

L LongfWdF r Dn.

UP Pitch 0

-5

Angular veloclty, rad /sec

Angle of attack, deg Nos:ok UP 0

240 250 2 -

I I I \+Break-away from tanker 3;;

Mach number

40 Up only

0

0 Aileron

14 Lt.


stick position, d y

0 StobilatorA 48

0 Aileron 1 OPushStabilator Ib

12

5

Rt Stick force,

"L Y 2 4 z UP

- 0 Pitch Dn. q - -2 Y 5 1 Rt. - 0 Roll = Lt.

-5

P P =


J Normal

load factor

1 I

Angular accel. rad/sec2

I

Sideslip angle, deg

114



r Cantrd

surface position,

0

Brake open

300 .6 200 I O 0

Brake open

300 .6 200 I O 0

Mach number

12 Right 40

0

0 Aileron

Dn spoiler Left alleron O

Rudder 0 Lt. -7 Control stick position,

Rear deg

Uponly posmon,

Lt 14

12 IO 8

TE. UP L Stabilatar 0 100 R t.

Rudder force, Ib 0

r Transverse o

L LanLwdg

o Stabilatw J 12

0 Aileron 1 Rt.

Stick

2 OpuShStablatar farce, Rt I b l 12

5

I

2

Load factor 2

UP

-.5 -2 r Dn.

Pitch 0 0 Pitch UP

Dn.

3 5 Angular R t Rt.

L t. Lt. -3 -5 .5 .8

velocity, Roll 0 0 Roll rad /sec

Yaw Rtb R t.

U. Lt. 0 Yaw

-.5 -.E


6

0

6

Nose It.

Nose rt. IO Nose

UP 0

0 IO 20 30 40 50 60 Time, sec.

_I

Normal load factor

I 1 Angular accet. rad/sec2

Sideslip angle, de9

Figure 62. - Data obtained during l e f t tu rn and landing. Airplane weight, 15,750 pounds; center of gravi ty a t 22.2 percent E ; f l i gh t 9.

NACA RM L57El7


0 I I I I I HP


600 - - - 500

1.2 -

.2

.4 -

.6 - - 300

.8 - - 400 1.0

- - - -

- -

200 Vi 1 -

- - IO0 M

12 Dn

40 Up only

0

0 Aileron aileron 0

surface Rudder O Dasitim. L t 14

Control Lt.

Rudder

r Load

factor L 7

10 8 12

TE.

Stabilator 0 UP

100 R t.

force, Ib 0

0 StabilatorJ 48

0 Aileron 1

Rear

Rt

Stick

2 OpuShStabilator Ibl

Mach number


i" Control stick position, de,g

Angular velocity, rad/sec


1 Rt

-rawverse 0 ' 5

2

L o n g Y g I

Pitch 0 2' Pitch UP

Dn.

2

-.5 3

Roll 0

Dn. -2 5

0 Lt, Roll R t.

L t.

R t.

-3 -5

.5 .8 Rt.

Lt. Yaw 0 0 Yaw

R t.

Lt -.5 -.8

6

0

6

Nose It.

Nose rt. 10 Nose

UP 0

0 IO 30 40 50 60 2o Time, sec.

Normal

factor load

1 Angular

I

accet. rad/sec 2

I

i Sideslip angle, deg

Figure 63.- Data obtained during landing approach, large directional oscillations. Airplane weight, 16,450 pounds; center of gravity at 22.5 percent ,E ; f l i gh t 26.

116 NACA RM L37E17



r Control

surface position,

L Rudder

r

' 0 1 0 m € 7 i i

factor

L L O " g ? j 2

[ Dn.

Angular R t.

L t.

Pitch 0

-.5 3

UP 0"' P i t h Dn.

-2 5 Rt.

velocity, Roll 0 O Lt. rod /sec

Roll

-3 -5 L YawRtO

.5 .8

0 Yaw R t.

u. Lt. -.5 -.8

6

0

6

Nose It.

Nose rt. IO

Angle of Nose

UP ottock, deg 0

I I I I

0 I

IO 20 30 40 50 60 Time, sec.

(a>

Figure 64. - Data obtained weight , 16,100 pounds ;

durira landing with large sideslip.

Normol load factor

1 I Angulor occet. rad/sec2

Sideslip angle, deg

Airplane center of gravi ty a t 23-2 percent C; f l i gh t 40.

-



r Control

surface position,

l0,OOO

600

400

12 Dn

Left oileron 0

Rudder 0 Lt

L Rudder

Load r factor 2 r "v

L Long?$ 'rnL

Pitch 0 - - ---..:' up' : q

Dn. 4 -.5

Angular I

velocity, rad /sec

I Roll 0

-3 Lt.

Yaw 0

-.5 Lt.


I 60 70

Mach number

U: Up only

Right

i" spoiler position,

Control stick

w d 48 R+

0 Stobi lotor1

0 Aileron 1 Stick

OpushStobilotor force'

12 - 5

. I

Ibl

- -

= 2

oUp Pitch f Dn.

x 5 -2

z Rt. - 0 Roll E Lt. =I -5

-

1

Normal load factor

1 1 Angular accet. rod/sec2

Sideslip angle, deg


118 NACA RM L57E17

c > L’

0 0 c

2 -0 U 0 -

IO

9

8

7

6

5

4

3

2

I

0

- I

-2

-3

-4

-50 100 200 300 400 500 600 700 Indicated Airspeed, Vi, knots

Figure 65.- Cornpaxison of measured normal load factors obtained during general flying with the V-n diagram. Pressure altitude, sea level t o 15,000 f ee t .

c >

0 L-

0 t

2

IO

9

8

7

6

5

4

3

2

I

0

- I

-2

-3

- 4

-50 .2 4 .6 .8 1.0 1.2 Mach number, M

Figure 66.- Cornpasison of measured normal load factors obtained during general flying with the M-n diagram. Pressure alt i tude, sea level t o 15,000 fee t .

120 FI NACA RM L57El7

> c c

Indicated Airspeed, Vi, knots

Figure 67.- Comparison of measured normal load factors obtained during general. flying with the service V-n diagram. Pressure altitude, l5,OOO f e e t t o 30,000 f ee t .

> c

t

Mach number, M

Figure 68.- Comparison of measured normal load factors obtained during general flying with the service M-n diagram. Pressure alt i tude, 15,000 f e e t t o 30,000 f ee t .

122

"0 100 200 300 400 500 Indicated Airspeed, Vi, knots.

Figure 69.- Comparison of measured normal load factors obtained during general flying with the service V-n diagram. Pressure altitude, 3O,OOO f e e t t o 42,000 fee t . -

> c

4 .6 8 I. 0 I. 2 Mach number, M

Figure 7.0.- Comparison of measured normal load factors obtained during general flying with the service M-n diagram. Pressure altitude, 30,000 f e e t t o 42,000 fee t .

124 I NACA RM L’j7E17

6

5

4

3

2

I

0

-I

-2

-3 I I I I 0 100 200 300 400 500 600

Indicated Airspeed, Vi, knots.

Figure 71.- Comparison of measured normal load factors obtained during LAIS missions with the service V-n diagram. Pressure altitude, sea l e v e l t o 13,000 f ee t . Airplane configuration, external stores carried.

I

- Z "0 .2 .4 .6 .8 I. 0

Mach number, M

Figure 72.- Comparison of measured normal load factors obtained during LABS missions with the service M-n diagram. Pressure altitude, sea

. l e v e l t o l5,OOO fee t . Airplane configuration, external stores carried.

126 - NACA RM ~ 3 7 ~ 1 7

> C


Figure 73.- Comparison of measured normal load factors obtained during LABS missions with the service V-n diagram. Pressure altitude, 17,000 f e e t t o 30,000 f ee t . Airplane configuration, external stores c a m ied.

6

5

4

c ’ 3

- I

-2

I I I I I I I I I I I I --

,/ and altitude of 15,000 feet. -- I weight of 18,500 pounds --

M-n diagram for airplane f

“0 .2 .4 .6 .8 I . 0 Mach number, M

Figure 74.- Comparison of measured normal load factors obtained during LABS missions with the service M-n diagram. Pressure altitude, 13,000 f e e t t o 30,000 f ee t . Airplane configuration, external stores carried.

-.i .

I -

128 I NACA RM ~ 3 7 ~ 1 7

Indicated Airspeed, Vi, knots

Figure 75.- Comparison of measured normal load factors obtained during dive bombing missions with the service V-n diagram. Pressure altitude, s ea l eve l t o 15,000 feet . Airplane configuration, external stores carried.

7 1 1 1 1 1 1 1 1 1 1 ~ ~ --

-- weight of 18,500 pounds 0

-r and altitude of 5,000 feet.

M-n diagram for airplane

‘Mach number, M

Figure 76.- Comparison of measured normal load factors obtained during dive bombing missions with the service M-n diagram. Pressure alt i tude, s ea l eve l t o 15,000 f ee t . Airplane configuration, external stores carried..

> c


0

Figure 77.- Comparison of measured normal load factors obtained during dive bombing missions with the service V-n diagram. Pressure altitude, 15,000 f e e t t o 30,000 f e e t . Airplane configuration, external stores carried.

> c

-U U 0

Mach number, M Figure 78.- Comparison of measured normal load factors obtained during

dive bombing missions with the service M-n diagram. Pressure altitude, 17,000 f e e t t o 30,000 feet. Airplane configuration, external stores carried .

V-n diagram for airplane weight of 18,500 pounds and 3,000 feet a l t i tude.

Service limit , clean

Service l i m i t , clean ', \

Design l i m i t

Indicated Airspeed, Vi, Knots

Figure 79.- Comparison of measured normal load factors obtained during general flying and bombing missions with the V-n diagram. Pressure alt i tude, sea l e v e l t o 42,000 f ee t .

133

Q

10

9

8

7

6

5

4

3

0 2

k 0

k rd

*

-Po

% rl

3 k 1 k2

0

-1

-2

-3

-4

M-n diagram for airplane weight of 18,500 pounds and 5,000 f ee t a l t i t ude /

-

-

” \

-5 0 .2 .4 .6 .8 1.0 1.2

Design limit -

I I I I I I I I 1 I

Mach number, M

Figure 80.- Comparison of measured normal load factors obtained during general flying and bombing missions with the M-n diagram. Pressure a l t i tude , sea l eve l to 42,000 fee t .

NACA - Langley Field, Vd.

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