naca report of the grumman f8f-1

7/28/2019 NACA report of the Grumman F8F-1

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NACA RM NO. ~8~27 ci illll~ll:rll:rl~~~~~~~luN x..ANATIONAL ADVI/ 3 1176014379748 RJQJJ-ICS;-,-,-. _IES-EQCHMEMORANDUM

for-theBureau of Aeronautics, Department of the NavyFIJIGEl!TMEAS- S OF THE LONGITUDllUL STABIIJTY AND CONTR

CaCTERISTICS OF THE GRUMMAN 8F-1 AIRPUNE -TED NO. NACA 2379

By Arthur Assadourian and John P. Reederi

This paper presents the results of flight tests to determine longitudinal stability and control characteristics and the stallingacteristics of the Grumman F8F-1 airplane. The lateral and directstabili ty-and control characteristi cs were reported in reference

With the normal center-of-gravity position, approximately 26cent mean aerodynamic chord, the airplane was unstable below the speed of 180 miles per hour and neutrally stable with st ick free 180 to 340 miles per hour, while with the stick fixed it was unstbelow 220 miles per hour, end neutrally stable above 220 miles pewhen in the normal rated-power clean condition. The airplane wain the gliding condition and unstable below the trim speed in theapproach condition, stick fixed and stick free. The instabili ty rated-power and power-approach configurations was considered accebecause of the light stick forces involved. The airplane was genstable in the rated-power clean condition, stick fixed and stick in accelerated turns for the range of conditions tested. Howevercase of the aft center-of-gravity position (27.4 percent M.A.C.) lower speeds, the elevator control force became zero at the higheaccelerations reached, a condition the pilot considered very ObJectionable-The elevator deflection available was always sufficient and the pthe elevator trimming tabs was adequate except that the airplane not be trimmed below 14.0 miles per hour in the landing condition. nose-up change in trim due to adding power in the landing configurwas considered large enough to be objectionable. The stalling ch


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NACA RM No. SL8H27referenCe 1. Airspeed was measured with a swiveling static head, one chord length ahead of and slightly below the right wing tip., shielded total-head tube, also mounted ahead of the wing tip. Thespeed system was calibrated for posit ion error by means of a trailinairspeed bomb.

Calibrated airspeed as used herein corresponds to the reading standard A-N airspeed meter connected to a pitot-static system thafree from position error and is defined by the formula

vc = 45 .08fo&where V, is in miles per hour, qc is the difference between totapressure and correct static pressure in inches of water, and f, compressibili ty correction factor at sea level (reference 2).

Control-stick forces were measured by means of a strain-gage ratus installed in a control column, replacing the service stick ahaving the same length, approximately 16.5 inches from the hinge the center of the grip. Sideslip angles were measured by a yaw vamounted one chord length ahead of the left wing tip. The sideslip were not corrected for any angularity in the flow existing at the vane. Previous tests made with similar type airplanes showed that correction would amount to no more than 2.

TEST RESULTS AND DISCUSSION .The results of the tests are evaluated in terms of the specifi-cations of reference 3.

Dynamic Longitudinal StabilityShort-period Oscillations were induced in the rated-power clecondition at several speeds for each of two center-of-gravity posiat an altitude of approximately 10,000 feet. The procedure used wtrim the airplane, then abruptly pull up to approximately 2g and rthe control column. Time histories of these pull-ups are presented

f i@?X 4. Very small oscillations of the elevator occurred in someases, particularly at low speeds, but were satisfactorily dampedperiod of the elevator oscillation was short enough so that the ai


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4 NACA RM No. Static Longitudinal Stability

The static longitudinal stability was measur ed at two center-ogravity positions, approximately 26 ma 22 percent M.A.C. (mean aeThe forward shift of the cenynamic chord) with landing gear up.of'gravity due to lowering the lanaing gear was approximatel y 0.7 M*A.C., whereas fuel consumption could cause forward shifts of as as 3 percent M.A.C. The weight of the airplane at take-off was ab9500 pounds. In the presentation of the data, account has been tathe effect of varying fuel loads on weight and center-of-gravitfpositioncFigures 5 to 7 contain plots of the variation of elevator angelevator control force, aa sideslip angle in straight flight agaicalibrated airspeed. For this series of tests, the pilot held theplane at essentially zero angle of bank and, therefore, laterally flight can be assumed. Plots of the variation of the elevator angnormal-force coefficient and the variation of elevator force dividethe impact pressure Fe/qc with normal-force coefficient are presin figure 8 for the three conditions tested. The stick-fixed and

free-neutral points were determined from the slopes of these curveneutral points for a given lift coefficient are defined as the cen"egravity positions at which the slopes - ma dEk/% are zero- aeN dCNdetermination of the neutral points in the rated-power clean, glidisnd power-appr oach conditions for several normal-force coefficients shown in figure 9. Figure 10 shows the variation of stick-fixed astick-free neutral points with normal-force coefficient.

Itshould be noted that the neutral points are accurately detonly when their locations are reasonabl y close to the range of cengravity locations tested. When the neutral points are far removedthe normal center-of-gravity limits of the airplane, however, theirexact location is of little practical signifi cance and they serve to show whether or not the airplane is stable rather than the degrstability. The degree of stability in these cases is better indicaby the curves of elevator stick force and elevator angle against aThe requirements of reference 3 state that with the center ofat its rearward limit, the variation of elevator angle with speed have a stable slope within the speed range specified for a given fcondition, and the variation of the elevator stick force with speebe such that push forces are required to increase speed from trim pull forces to decrease speed. Information received from the Grum


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NACA RM No. SL8H27. . .. .... . . .I . . The following conclusions were reached regarding the static lo&I . .. . tudinal stability of the F8F-1 airplane:

:.I l .I. .1.: . . 1. Rated-power clean condition-- The airplane was unstable up . . approximately 220 miles per hour and neutrally stable from 220 , . :, . .. . -to 400 miles per hour, stick fixed, %nd unstable up to approximately180 miles per hour and neutral,ly stable from 180 to 340 miles per hstick free, for the aft center-of-gravity position (27.1 percent M.when trimmed for laterally level straight flight at 180 miles per hFor the forward center-of-gravity kosition (22.1 percent M.A*C.), thairplane was unstable below the trim speed and stable above, stick and stick free.

2. Gliding condition.- The airplane was stable, stick fixed anstick free, for the aft center-of-gravity position (25.7 percent M.and the fo?Xmd center-of-gravity position (22.2 percent M.A.C.) whtrimmed for laterally level straight flight at 14.0 miles per hour.3. Power-approach donditionL- The airplane was unstable below

trim speed of 110 miles per hour and stable above stick fixed and free, for the aft center-of-gravity position (24.6 percent M.A.C.).the, forward center-of-gravJty position (21.2 percent M.A.C.) the airwas unstable below 100 miles per hour and stable above, sti ck fixed stick free04. General.- The static-longitudinal-stability requirements ofreference 3 were satisfied only fop the airplane in the gliding conand above the trim speed'for the power-approach condition. The inst

bility-of the airplane in the rated-power and power-approach conditiowas considered mild by the,pilot a& acceptable because of the lightstick forces.The effects of power on the stability of this ai rplane a&low were rather large. These power effects were, however, less than mibe expected for an airplane with an engine and propeller of such larsize in relation to the rest of the airplane. The power effects weprobably kept at a minimum due to the fact that the thrust axis was

tilted down 3-60 from a reference line parallel to the airplane whewing is at the angle of zero lift. ,Longitudinal Control in Accelertited Flight

The longitudinal stability end control characteristics in accel


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6 NACA HM No. SThe stick-fixed maneuver points were determined for a normal-forccoefficient at the middle of the range covered for each speed as thecenter-of-gravity position where values of the slope a6,/acN are

in figure 13. The stick-fixed qeuver pointwas at approximately cent M.A.C. at an indicated airspeed of about 350 miles per hour annormal-force coefficient of 0.3 and moved aft with decreasing speed increasing normal-force coefficient.

The stick-free maneuver points were also determined from figure for an acceleration at the middle of the range covered at each speedstick-free maneuver points are the center-of-mvity.positions wherevalues of the slope dFe/dn are zero in figure 13. The symbol nrepresents normal acceleration in gravitational units. The stick-freemaneuver point for a 3g turn at 350 miles per hour was at approxi-mately 28 percent M.A.C. end moved aft as the speed and accelerationdecreased.Throughout the test range of normal-force coefficients and acceations in right and left turns, the airplane was generally stable, fixed and stick free, for both forward snd aft center-of-gravity postions,'the stability being more positive for the Forward center-of-gravity position, the higher speed, and right turns.However, in the case of the aft center-of-gravity position (27.cent M-A-C.) and the lower speed, the elevator control force becameat the highest accelerations reached (fig. 11). The pilot consideredthis condition very objectionable and the requirement of reference which states that the elevator control-force,gradient in steady turnshall never be less than 3 pounds per g was not satisfied.Longitudinal control in landing.- The longitudinal control inlanding was considered satisfactory. Figure 14 presents a time histoof a typical landing from which it will be noted that with the centeof gravity at approximately 24 percent M.A.C*, landing gear down, a16O up-elevator or about two-thirds of that available wss used to laAnother landing record showed that, with the center of gravity at a20 percent, approximately 21 up-elevator, or 3O less than that avaiable, was required to land. Although in both cases it was not possito trim the airl,lane below approximately 140 miles per hour, the ele

control forces were well below the maximum of 35 pounds considered factory by the standards of reference 3-Longitudinal control in take-offs.- The power of the elevator control the longitudinal attitude of the airplane during take-offs found to be adequate. Time histories of a few take-offs were presenin reference 1.


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NACA HM No. SL8E27. . . .. l. :. .. . . ... ::. .. . .. . .:.: . .. . ... '.:. .

conditions. It was possible to trim the elevator control forces throughout the test center-of-gravity range from the highest speereached to within a few percent of the stalling speed for all but *ding condition. The minimum trim speed for this condition was140 miles per hour.

Trim changes due to flaps and power.- The longitudinal trim due to flaps, landing gear, and power were measured with the centegravity at approximately 26 percent M.A.C. (landing gear down) anthe elevator trimming tab set at- 21.2o (full) nose-up at 140 mileshour. The airplane was trimmed with flaps and landing gear do-urnidling, and successive changes in configuration were made as showthe table.Position of- Approximate eleLanding Power setting , control forcFlaps gear (1%)

Down Down ' Engine idling 0 Down Down Normal rated (41.5 in. Hgat 2600 rpm) 24.2 pusDown b Normal rated (41.5 in. Hgat 2600 rpm) 29.2 pus

VP VP Normal rated (41.5 in.- Hgat 2600 z-pm) 1 43.0 pus

The large trim change, requiring an elevator control push for43 pounds, n going from the landing to the normal rated-power clecondition did not satisfy the standards of reference 3 which stateit should be possible to maintain a given trim speed using any conation of engine power, flaps, or gear setting without exerting ppull forces greater than 35 pounds. In particular, the nose-up cin trim due to adding power was considered large enough to be objeable. Since the pilot observed that the trim change in going fromlanding to the normal rated-power clean condition was by far the no other configuration was tested.


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induced b'y rudder kicks are presented in figure 16, from which tydata the summary plot of figure 17 was obtained. In'this figure, in normal acceleration is plotted against change in rudder angle. should be noted that the normal-acceleration changes which occur dependent upon the time interval the rudder is held in the deflectposition and, therefore, the values presented in figure 17 shouldconsidered as only qualitative. The solid curves represent the facceleration peak after the rudder is kicked. In left rudder kicairplane initially pitched up but then pitched down more violently.dashed portion of the curves represent this second peak. The pitcmotion is caused by the combined effect of the gyroscopic action propeller and the pitching moment due to sideslip. In the case orudder kicks, both these effects tend to pitch the airplane downleft rudder kicks, the airplane initially pitches up because of tgyroscopic moments, but then pitches down as the sideslip builds Figure 18 presents typical time histories of maneuvers in which tstick WEEI eleased from steady sideslips with the rudder fixed= results, therefore, isolate the effect of pitching moment due to slip. It is seen from this figure that the acceleration increasedthe center of gravity was moved aft. From data of this type, theplot of figure 19, presenting change in normal acceleration againslip angle, was obtained. The large changes in normal accelerationto release of the stick were considered unsatisfactory.

The results of these tests are to be cbnsidered as a preliminainvestigation towards formulating some flying-qualities requiremenlimiting the amount of pitching motion of aircraft induced by yawmotion. As the results show, the elevator forces required in stesideslips, though they do not appear excessive, may, if coupled wlow stick force per g gradient, cause high accelerations. Therefoairplanes having low stick force per g gradients should have corrspondingly small elevator stick forces due to sideslip. .

Hinge moments.- The elevator hinge-moment parameters % awere analytically determined using the data obtained in acceleratedturns* A correction to account for the moment introduced by the weight was applied to the data. The values of the rate of changeelevator hinge moment with angle of attack (2% and with elevatordeflection Chge were -0.0014 and -0.0068 per degree, respectively.

CONTROLFRICTION


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NACA RM No* SL8H27

Control

ElevatorRudderAileron

Friction at neutraldeflection(lb)fl$f6+124

Maximum allowable frictiat neutral deflection(lb)+3t7+2

STALIXNG CRARACTERISTICSTime histories of stall approaches and stalls in the rated-pclean, gliding, and power-approach conditions are given in figureThe stalling characteristics of the F8F-1 airplane for various crations were as follows: i1. Rated-power clean (fig. 20(a)).- Stalls in this condition somewhat in nature and speed, probably due to differences in sidangle and to rate of change of angle of attack during different In a typical case, however, there was a mild roll to the left wbe easily corrected followed by an abrupt roll to the right. Asteep nose-up attitude gave some stall warning in this condition.stalling characteristics were considered satisfactory because ofrelatively small initial rolling velocity.2. Gliding (fig* 20(b)).- The elevator control force was incorrect direction but was light. At the stall, the nose droppedwith a sharp roll to the left. The stalling characteristics in condition were acceptable.3. Power-approach (fig. 20(c)).- Stalls in this condition wcharacterized by a mild roll to the left accompanied by mild bufThe stalling characteristics were considered acceptable be'cause mildness.4. Landing.- Rearward stick movement accompanied by a lighteof the elevator stick forces provided a good stall warning as that 94 miles per hour was approached. At the stall, the nose droand the airplane rolled abruptly to the right about 15'. Stalling


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10 NACAnose-down pitching characterized the stall- Stall warningsatisfactory because of the rudder buffeting and the steep

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stall was consfdered unsatisfactory because of the severe pitching.The stalling characteristics were considered acceptable for afigurations, even though definite buffeting prior to the stall wasreadily discernible by the pilot in most cases due to the general 6f the airplane at the lower speeds. This shaking could not be coered a good stall warnFng because it covered too large a speed ranfrom about 115 miles per hour to the stall. Normal stall recovery cedure was used to regain control of the airplane, correcting for roll at the same time.A time history of a 3g wind-up turn to the stall in the rated-clean condition is shown in figure 21. No warning preceded the stwhich was characterized by an abrupt reduction in acceleratXon andpitching velocity but accompanied by no roll.

CONCLUSIONS 0The conclusions reached regarding the longitudinal stability control characteristics of the Grumman F8F-1 airplane (BuAer No. 9maybe summar zed as follows:1. Small-amplitude, short-period longitudinal oscillations prby abruptly deflecting and releasing the elevator were satisfactorilydamped.2. In the rated-power clean condition for the range of center-

gravity positions tested (22.1 to 2'7.1 percent M.A.C.), the airplanewas unstable, stick fixed and stick free, below the trim speed of180 miles per hour. The airplane was stable, however, above the tspeed for the forward center-of-gravity position (22.1 percent M*Astick fixed and free.3. Throughout the test center-of-gravity range, both stick fixand stick free, the airplane was stable in the gliding condition aunstable below the trim speed in the power-approach condition.4. The instability in.the rated-power and power-approach confirations was considered acceptable because of the light stick forcesinvolved.5* The airplane was generally stable in the rated-power clean


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NACA RM No. SL8H276. There was always sufficient elevator deflection.available longitudinal control during take-off and landing or to reach the in straight or+.zrn.ing flight.7. The power of the elevator triting tabs was adequate exckhat the airplane could not be trimmed below approximately 140 mhour in the landkng configuration. The trim change in going fromlanding to the normal rated-power clean condition was in excess standards of reference 3. In particular, the nose-up change in

due to adding power was large enough to be considered objectionable8. Although no definite buffeting was present except in the off condition, the stalling characteristics were considered accepand the recovery was normal.9. Changes 2t-1 ongitudinal trim in the form of normal acceleand pitching induced by yawing velocity and sideslip were consideobjectionable and it is proposed that flying-qualities requirement

should contain a provision to limit the change. *

Langley Aeronautical LaboratoryNational Advisory Committee for AeronauticsLangley Field, Va. '

L Arthur AssadourianAeronautical Research S

4John P. ReederAeronautical Research SApproved: Melvin N. CoughChief of Flight Research DivisionBKB


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REFEZENCB

1. Crane, H. L., and Reeder, J. P.: Flight Measurements of LaterDirectional Stability and Control Characteristics of theGrumman F8F-1 Airplane - TEZI No. NACA 2379. NACA RM No. SBur. Aero., 1947*2. Aiken, Wil liam S., Jr.: Standard Nomenclature for Airspeeds Tables and Charts for Use in Calculation of Airspeed. NAC'TN No. 1120, 1946.3. Anon: Specification for Stability and Control Characteristics Airplanes. SR-llgA, Bur. Aero., April 7, 1945.

NACA RM NO. ~~8~27


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1 . ..

: TABLE I.,... . .. .1.

. . GENERALSPECIFICATIONS OF TEE AIRl?L.@E. -0. .. . .. . . :. . Make and designation . . . . . . . . . . . . . . . . . . . Grumman. . e (Btier No. Engine . , . . . . . . . . . . . . . . . . Pratt and Whitney R-28ODoublePower ratingsTake-off . . . . . . . . . . . . . 2100 hp at 2800 rpm at sea

Mili tary . . . . . . . . . . . . .Normal maximum (low blower) .* . . . 1600 hp at 2800 rpm at 16,01703 hp at 2600 rpm at 7(high blower). . . . 1450 hp at 2600 rpm at 18,5Propeller ............. Hydraulically-Controlled Four-Constant-Speed - AeModel ........................... A 6Bladenumber ......................Basic pitch settings .............. Max 63.0, Min

Diameter ........................Fuel capacity, galMaintank ........................Droppable (belly) .................... 100 Droppable (wings) ........... . ........Cil capacity, galOne tank (in engine compartment) . . . . . . . . . . . . . . .

. War emergency power system fluid, galOne tank (in engine compartment) . . . . . . . . . . . . . . .GeneralSpan (wings spread), ft . . . . . . . . . . . . . . . . . .Span (wings folded), ft . . . . . . . . . . . . . . . . . .Length (over all), ft . . . . . . . . . . . . . . . . . . .Length (tail wheel on ground), ft ., . . . . . . . . . . . . Length (tail wheel on ground), propeller blade vertical ftWeight for tests, approximate, lb . . . . . . . . . . &GO to -_-_ _ _-..Jypc.+


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14 NACA RM'N. . . .. .: . : GENERALSPECIFICATIONS OF THE AIRPLANE - Concluded. . . .. . .. . .. .. . .

:: l . .Wing8.. . . . Area,sqft.......................... .. . : Airfoil section. . .. . Root l . . l l . . . . . . . . . . . . . . . . . . modifiedTip ...........................Meanaerodynamic chord, in. .................Leading-edge M.A.C. aft of leading edge of root chord, in. .Root chord, in. ........................

Tip chord (6 in. inboard of actual tip) in. .........Incidence, deg ........................Dihedral ..........................Sweepback of leading edge, deg ...............wing flapsArea, total, sqft .....................Deflection, maximum down, deg ................AileronsArea, total, sq ft .....................Spring-tab area, total, sq ft ................Trimming-tab area, sq ft ..................Trimming-tab deflection angle, deg .............Horizontal tailSpan,in ...........................Total area, sqft ......................

Elevator area (including tabs) sq ft .............Elevator trinrming-tab area (total) sq ft ..........Elevator tab range, deg (approx.) ...........Tail incidence, deg 8 up,. / ..................Vertical tail (configuration 3 of reierence 1)Totalarea, sq ft. .....................Rudder area, sqft .....................Rudder-tab area, sq ft ...................

Finoffset,deg. ......................Rudder-tab range, deg (approx.) ............... c---. . - - t$c&&-

--- .


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NACA RM ND. SL8H27

=jq$; Je4 5Verfical tail, Z/inches above fuselage referen

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Horizontal fail / LFinches from center

Winy and dlerm, new in board emf of aileroTFigure 2.- Section views through the aerodynamic surfaces of the Gr


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Figure 3.- Linkage between the control stick and elevator of the Grumman F8F-1 air

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Time, set(a) J50 mph. (6) 250 mph.

(a) Center of gravity at 25.2 percent mean aerodynamic Figure 4.- Short-period longitudinal oscillations in the rated-pawcondition, Grumman F8F -1.

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. . . .. . .: . . NACA RM No. SL8H27. . . .. .: . .. .. . .. . .. . .. . . .. . .. *; .:.

BO 2 4 6 0 2 4J/me, set(6) 250 mph.

.I (D) Center af gravity at 20.7 percent mean aerodynamic

. . . .NACA RM NZJ. SL8K27


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I20 I60 zoo 240 280 320 360Calibrated ahpeed , mph

(a) Center of gravity at 27.1 percent mean aerodynamic chord elevator trimming tabs set at neutral.Figure 5.- Static longitudinal stability characteristics of the Grum


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; ;. ..:.a

30 120 160 200 240 260 320 360 40Cali+7ted c7/rspee'd, mph

(b) Center of gravity at 22.1 ,percent mean aerodynamic chord and elevator trimming tabs set 2 nose down,Figure 5. - Concluded.

. . .. . . .


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Calibrated airspeed I mph(a) Center 2f gravity at 25.7 percent mean aerodynamic chord


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!(b) Center of gravity at 21.2 mean aerodynkmi c chord and elevator trimming tabs set at full nose. up (not trim ).

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Figure 7.- Concluded.

. . .. .:.. .


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4do 240 do do do do ho sb 8&- 4do do hi0 do do 40 &?Qb Cdibfuted ahpeed f mph *9 /Figure 10. - Summary plots of the variation of the neutral points with normal-force coefficient,Grumman F8F-1 airplane,, N. *

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60oRight furns0 Leff fums

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Change in nor-m a/ accelevu fiun J9(a) 200 mph, (b) 350 mph, ;;d v

Figure 11. - Variationof elevator control force with change in normal acceleration for constaairspeed turns in the rated-power clean condition, Grumman F8F-1 airplane.

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Figure 12. - Variation of elevator angle with normal-force cogfficient for constant airspeed turns inthe rated -power clean condition, Grumman F8F -1 airplane.

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