Document
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Copy No. /
RM No. SL7L31
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NACA
Kiu'$ -J FILE
RESEARCH MEMORANDUM
for the
Bureau of Aeronautics, Navy Department
FLIGET MEASUREMENTS OF LATERAL AND DIRECTIONAL
STABILITY AND CONTROL CHARACTERISTICS
OF THE GRUMMAN F8F-1 Anw.Am
TED NO. NACA By
H. L. Crane and J. P. Reeder
Langley Memorial Aeronautical Laboratory
Langley Field, Va.
JN CANC-61,
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RESTRICTED for Awww* ^ x NAC ^ RM No. L I L31 S V eI A=SORY COMMr__= 1 FOR A-.EPONAUTICS NATION AL RESE ARCH MEMCR WDUM for the Bureau of Aerona.utice, Navy Department TF.R AL A ND DIRECTION AI, FLIGHT MEASURE ENTS OF L A STABILITY AND CONTROL CH.kRACTERISTICS OF = GPUMM A N F8F-1 A.IRFLA.NE
TED NO. NA.0 2?79
By H. L. Crane and J. ' 0 . Peeler S U M M A R Y Th'_ g naoer presents the resl:lts of flight tests to determine the lateral and directional stability and control characte r istics of the Grumman F 5F-1 airplane with three vertical-tail confimzrations. The -1 to Dresente^i herein have no bearin g on the performance characterist'cs of the 9{rDlane, which were not measured but which were considered to be excep- tionally good. The conclusions re9.ched regardinAz the l .atera' And direc- tional stability and control characteristics may be summarized 9.s follows: 1. It was found that the directional stability was Door with the production vertical tail. Addition of a 12-inch extension to the vertical fin and rudder produced a desirab l e improvement in directional stability and control chara.cter'stics. However, farther enlarwement of the vertical tail would be renuired to make the direct i ontl stability satisfactory in P-11 respects.
2. There was a tendency for the ruller control force to overbalance at large an71es of right sidesli p with the mod'fied vertical trills. There was no such tend-enc7 with the production tail configuration which included a lors.a.l fin. It was concluded that the dorsal f'n should have been retained on the modified vertical tails.
3. The aileron control ch^racteri g t'_cs ^:.rere better than those of many cemoarable a.irnlanes which have been tested. However, the ailerons did riot satisfy the Navy re q uirements for s?.t.isfactory fla r ing qualities w'th regard inn, effectiveness.
to e?ther control forces or roll 4. The p ower of the rudder triamina tab p roved to be inadeaulte and the tab should be enlarged And/or be provided. with ?n increased leflectlon range.
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NACA RM No. L7L31 li.
I N T R O D U C T I O N At the request of the Bureau of Aeronautics, Navy De partment, flight tests have been made to determine the flying qualities of the Grumman F8F-1 airplanes (BuAer nos. 94873 and 90 1 +61) with three vertical-tail ^.
configurations. Flight tests were made with XFBF-1 airplane 90461 between May 10 and August 4, 1945 when the landing gear and propeller were damaged • during an emergency landing. The test program was resumed unler lower priority on June 6, 1946 when a second airplane, F8F-1 airplane 94873, was provided by the Bureau of Aeronautics. The results of the tests to deter- mine the lateral and directional stability and control characteristics are presented in this report. The investigation included a rather complete measi:rement of the lateral and directional stability and control charac- teristics with the production FBF-1 vertical tail plus enough supple- mentary tests with the modified vertical tails to determine which config- uration was the most satisfactory.
DE SCR IPT TON OF AIRPLANE
The F8F-1 airplane was a single-engine, single-place, low-wing, shipboard fighter. The F8F-1 airplane was equipped with slotted flaps a_-rd a conventional-type retractable landing gear. The construction was all metal except for the fabric covering on the control surfaces. Tho Frise ailerons were equipped with spring tabs. The rudder was horn- balanced and the elevator had an cverhanging balance. Test airplane 90461 was equipped with trimming tabs on the rudder and elevator but had none "on the ailerons. This airplane was equipped with dorsal and ventral fins.
The fin was offset 2 0 leading edge left. During a large portion of the tests a 150-gallon external fuel tank was mounted below the fuselage.
There were no dive-recovery flaps on airplane 00461. Power was supplied by a Pratt & Whitney R-2800-34 engine which turned a four-blade Pero Products propeller. The take-off weight of the airplane was approximately 0000 pounds with no external stores or 10,000 pounds with full belly tank.
Photographs of the F8F-1 airplane with and without the 150-gallon belly tank are presented in figure 1. A-three-view drawing, cross sections of the wing and aileron, and cross sections of the horizontal and vertical tail are shown in figure 2. General specificat_ons of the airplane are given in table I.
The Grumman F ,9F-1 airplane 94873 differed from airplane 00461 in that it had a fin setting of 1.50 J. leading edge left, and was equipped with dive-recovery flaps and an aileron trimming tab. .Airplane 94873 was used to test the modified vertical-tail configurations.
i NACA RM No. L7L31 Figure 3 presents the variation o'' rudder deflection with right- rudder-pedal position. Rudder deflection was measured with respect to ^:^^ The friction of the rudder system amounted to ±6 pounds of the fin.
' pedal force. Figure 4 presents the aileron and spring-tab linkage characteristics. The relation between ailer'n deflection and control- stick position with no load on the system is shown in figure 4(a) (aileron tabs locked). The variation of aileron-spring-tab deflection with control-stick position Is shown in figure 4(b) (ailerons locked).
Figures 4(c) and 4(d) present the variaticn of aileron-spring-tab deflection with stick force for the left and right tab, respectively.
The variation of tab deflection with stick force for the left tab was obtained with the left aileron locked and the right aileron free in order to measure only the resistance of the left tab. The process was then tab.
reversed for the right The aileron control system included stops on the tabs, but there were no stops at the at the stick and stops In this paper the direction of tab deflection is given with ailerons.
respect to the control surface.
Sketches of the three tail configurations are shown in figure 5(R).
The F8F-1 production tail, vertical-t ail configuration 1, was equipped with a dorsal fin and the rudder gap was not sealed. The two modified vertical-tail configurations were installed by the Grumman Company.
The first modified configuration, hereinafter designated as tRil config- uration 2, resulted from adding 12 inches to the span of the fin by Inserting an extension between the fuselage and the production fin and adding a 2-inch trailing-edge strip, which is shown in figure 5(b), to the rudder and trimming tab. The second modified configuration, herein- after designated as tail con f iguration 3, had both the fin and rudder span extended 12 inches. The 2-inch trailing-edge strip was retained only on the trimming tab. Tail configurations 2 and 3 did not include a dorsal fin,but had the rudder gap sealed.
The rudder control system of F8F-1 airplane 94973 included a spring on the left rudder cable which had been installed to reduce the tendency of the rudder forces to lighten in right sideslip with tail config- urations 2 and 3. The pedal-force gradient produced by the spring, which is illustrated in figure 6, was approximately 2.6 pounds per degree of left rudder deflection.
NACA RM No.
4 L7,31
I N S T R U M F N T A T I 0 N The following instruments were mounted in the airplane: NACA instrument Measures. quantity Timer(synchronizing, all records) Time Airspeed recorder and sensitive indicator Airspeed Control positions Control-position recorders Strain-gage pedal-force and stick-force Control forces recorders Sideslip-angle recorder and indicator Sideslip angle Normal, longitudinal, Three-component recording accelerometer, and transverse sensitive normal accelerometer, and indicating normal accelerometer accelerations (and angle of bank) ptl °.r velocit'es Gyroscopic rolling, pitching, and Ang yawing velocity recorders Flap position Position recorder Free-air temperature Flectrical resistance-bulb type thermometer Airspeed was measured with a swiveling static head, mounted 1 chord length ahead of and slightly below the right wind{ tip, and a shielded total head tube also mounted aheal. of the wing tip. The standard airspeed Indicator was re placed with a sensitive meter which was connected to the NACA. airspeed installation to enable the pilot to hold constant speed during maneuvers in which the Angle of sideslip varied. The airspeed position error with a trailing airspeed head.
system was calibrated for Calibrated airspeed as used herein corresponds to the reading of a standars A-N g jrspeed meter connecter to a pitot-static system that is free from position error end is defined by the formula
V = 45.08 f
c where V c in miles Der hour Q difference between total pressure and correct static Dressure in inches of water c f o compressibility correction factor at sea level NACA RM No. L7L31 5 Control positions were measured with both mechanical and electrical . recorders. Transmitting elements of electrical recorders were mounted on the control surfaces, aileron tabs, and the wing flaps. A mechanical control-position recorder was connected into the rudder and the elevator ..
system near the control stick. A chain attached to the control stick to make it possible for the pilot to hold constant stick deflection also served as a check.on the aileron-control-position recorder.
In order to measure control forces the service stick was replaced with one of the same length, approximately 16.5 inches from hinge line to the center of the grip, which contained a strain-gage installation for stick-force measurement. The rudder pedals were modified to accom- modate a strain-gage installation.
TESTS, RESULTS, AND DISCUSSION The results of the tests are evaluated in terms of reference 1.
DYNAMIC LATERAL AND DIRECTIONAL STABILITY Control-free lateral oscillations were made to determine the effect of speed, altitude, of the belly tank, and vertical-tail configuration.
These tests were made in the clean condition with power for level flight either by abrupt deflection and freeing of the rudder or by releasing the controls in steady sideslips. Figures 7 and present time histories of typical oscillations of each type. Surmnary plots which show the period of the lateral oscillations and the number of cycles required to damp to one-half amplitude as a function of speed are shown in figure Q.
The data of figure 9 show that in all cases the oscillation damped to one-half amplitude in well under 1 cycle. The tests showed that the belly tank had little effect on the damping of the lateral oscilla- tions. The damping was increased by decreasing speed or altitude and by the increased size of vertical —tail configuration 3.
The period of the lateral oscillation was decreased noticeably with
tail configuration 3. The effects of altitude or of the belly tank on
the period were small. Shown in figure 9. the period ranged from approxi- mately 2 seconds at 350 miles per hour to 4.5 seconds at 150 miles per hour.
Although the damping of the lateral oscillation satisfied the renuirements of reference 1, a persistent small-amplitude oscillation of higher fre q uency, which was annoying to the pilot, often occurred even in smooth air. Time histories of such oscillations are presented in figure 10. The oscillation shown in figure 10(a) was induced by a NACA FM No. L7L31 rudder kick and was obtained with an empty belly tank installed and with the fvsela,,^e tank approximately half full (at a speed of 250 miles per hour). In some cases the oscillation was noted mainly in terms of ' ^•^ lateral acceleration (fig. 10(a)) and in other cases the oscillation • occurred in the three measured components of acceleration (fig. 10(b)) Variations of propeller pitch and fuel sloshing were considered as possible causes of the oscillation. The first possibility was eliminated by obtaining the oscillation with the propeller set against the low-pitch sto p s. It was then found that the oscillation could be induced by a maneuver such as an abrupt turn entry. During the oscillation there was a fluctration of the fuel gage which led to the conclusion that the oscillation was due to fuel sloshing.
The data of figu res 7 and R show that there was no tendency for the rudder itself to oscillate. The tendency of the spring-tab ailerons to oscillate was investtgated by abruptly deflecting and freeing the ailerons. Time histories of the maneuver are presented in fi gure 11.
Oscillat'on of the ailerons was comp letely damped in 12 cycles which satisfies the requirements of reference 1.
STATIC L_=AL AND DIRECTIONAL STABILITY Sideslip Due to Deflection of Ailerons; Rudder to Overcome ldverse Aileron Yaw The sideslip due to deflection of the ailerons with rudder fixed and the rudder re q uired to overcome adverse aileron yaw were measured. in
rolls out of 45 0
banked turns. Figure 12 contains time histories of rolls out of turns which were made by applying approximately two-thirds aileron deflection at 5000 feet in the clean condition with bell y tank off using power for level flight at.135 miles per hour. In these rolls the pilot attem.Dtel to coordinate the rudder and ailerons to maintain zero Sideslip. Similar roll-outs made with full aileron and rudder deflection showed. that there was a little more than sufficient rudder control to over- come the yawing moment due to ailerons. The rudder control force did not exceed 180 Dounls during these maneuvers.
The maximum change in sideslip angle, which was measured with the belly tank installed. and w'th no belly tank in rudder-fixed roll-outs at 135 miles per hour, is plotted. in figure 13 as a function of the change in total aileron anr yle. The effect of the belly tank on the amount of sideslip obtained was small. Figure 13 also presents data which show a reduction in sideslip angle of a pproximately 15 Dercent with the enlarged vertical tail of configurat'on 2 or 3. ComDarison of the d,.ta of figure 13 with the re q uirements of reference 1 indicates a lack of directional stability at moderate sideslip angles with the present aileron system. The deficiency in directional stability NAC A RM No. L7L 31 7 would become considerably greater if the aileron control system was modified to provide a maximum rolling effectiveness pb/2V of 0.00, the value required by reference 1.
.; Rudder-fixed ro'_ls out of ?g turns at an indicated airspeed of i 290 miles Der hour at 5000 feet were made in the rated power, clean • • condition with tail configuration 1. The data obtained are plotted in figure 14. In these maneuvers the Dilot attempted to maintain a constant acceleration as the airplane rolled. The pitchin g moments caused by sideslip and by yawing velocity during the roll made it difficult for the pilot to maintain exactly constant acceleration, and as a result the data obtained show some scatter. Figure 14(a) contains a plot, of the change in sideslip angle as a function of the change in total aileron Angle. In figure 14(b) the same data are plotted in another form as the variation of AB/CN, the change in sideslip angle divided by the normal-force coefficient, with pb/2V. Plotting in this form makes possible a comparison of the directional stability of different airplanes without regard tc the effect of normal-force coefficient on the change in sideslip angle or to relative aileron size.
The data were obtained in rolls out of 3g turns rather than in rolls out of more highly accelerated turns in order to limit the sidesli p angles obtained to safe values. It should be noted, however, that the yawing moments acting on the airplane will increase in proportion to the lift coefficient and as a result larger sideslip angles could be obtained in rolls out of turns at high accelerations.
The yawing moments also increase in proportion to pb /2V. If the aileron power were increased to provide a value of pb/2V of 0.09, the maximum sideslip angle would also be increased. Extrapolation of the available data on the assumption that the yawing moment varied linearly with sideslip angle indicates that a sideslip angle of 400 6g turn with a would result in a roll at 290 miles per hour from a pb/2V of 0.09. Actually, the directional stability probably increases at such large sideslip angles and the resultant angle of sideslip would be somewhat less than 40 0 . It appears, however, that the directional stability of the XFRF-1 airplane with tail configuration 1 is inade- quate to limit the sideslip angles in rolling maneuvers to reasonable values.
All measurements of directional stability with tail configura- tions 2 and 3 were made with a belly tank installed, and therefore no accelerated maneuvers were performed. The directional stability in Accelerated rolling maneuvers would be considerably improved with tail configurations 2 and but not enough to eliminate the.possibility of overloading the vertical tail. Modification to improve the lateral control system would increase the possibility of overloading the tail.
NACA RM No. L7L31 of Sideslip Characteristics I.:..
The sideslip characteristics were investigated in steady side- • slips with all three vertical-tail configurations. Tests were made with tail configuration 1 with and without the belly tank at an alti- tude of 5000 feet. Tests were also mode at 300 miles per hour in the rated power, clean condition at an altitude of 20,000 feet with tail ••^ configuration 1 wi sh the belly tank installed. Measurements of steady sidesli p characteristics were made at 5000 feet with belly tank installed with tail configurations 2 and 3. The test conditions and speeds were as follows: NAC ' RM No. L7L31 • pp O b • r^ ri • O_ 4-) 0^ 0- RV H O U r-1 r-1 r--1 rl H r) rl .-1 ri r1 ..
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NACA RM No. L7L31 The sideslip data which gave mensuremente of directional stability •^•• (variation of rudder angle and force with sideslip angle), dihedral effect (variation of aileron angle and force with sideslip angle), force •••, pitching moment due to sideslip (variation of elevator angle and with sideslip angle), and the side-force characteristics (variation of angle of bank with angle of sideslip) are shown in figures 15 to 22 for tail configuration 1.
The results of the tests with tail configuration 1 will be discussed first.
Directional stability.- The curves of rudder position against sideslip angle indicate that without the belly tank the rudder-fixed directional stability was always positive,but was low at small angles of sideslip. In the rated-power, clean condition near zero sideslip the values' of the variation of rudder angle with sideslip d8r/ds Increased from about 0.15 at 150 miles per hour to 0.5 at 250 miles per hour and decreased to 0.35 at 400 miles per hour. The variation of rudder force with sideslip angle indicated that there was a very slight degree of rudder-free directional stability in left sideslips at 150 miles per hour. At higher speeds the variation of rudder force with angle of sideslip was stable but relatively small.
Addition of the belly tank reduced the directional stability in all conditions so that the value of d5r /dp in the rated-power, clean condition increased from 0.10 at small anizles of left sideslip at 150 miles per hour to about 0.35 at 350 miles per hour and decreased to 0.25 at 305 miles per hour. The rudder-free directional stability was neutral in left sideslips up to 20 0 at 150 miles per hour. At higher speeds the variation of rudder force with sideslip angle was small in comparison with that of other fighter-type airplanes previously tested.
For example, the sideslip obtained per pound of pedal force was almost three times that obtained with the F6F-3 airplane at 300 miles per hour.
As a general rule, decreasing the power or lowering the flaps increased the directional stability. The effect of altitude on the directional stability was determined b y comparison of the data of figure 22 which were obtained AT, an altitude of 20,000 with those of figures 16(b) and 16(c) for an altitude of 5000 feet. The :;maparisun snowed that V redo for small angles of sideslip was reduced from approximately 0.25 at 5000 feet to 0.1 at 20,000 feet (300 mph, tank on).
Dihedral effect.- The effective dihedral was positive control-fixed in all test conditions with or without the belly tank. There was a slight increase in effective dihedral due to the belly tank. The control- -free effective dihedral was positive in all test conditions except at 90 miles per hour in the power-approach condition. In that condition the control-free effective dihedral was neutral with the belly tank off and very slightly positive with the belly tank on.
NACA RM No. L7L31 Pitchin.7 moment due to sideslip.- The variation of elevator position with angle of sideslip was usually not large. The variation ..
of elevator force with angle of sideslip is not presented in all cases because the elevator-force recorder used during the tests of tail config- uration 1 was found to be unreliable. However the data obtained with tail configuration 2 (fig. 23) indicate that a fairly large pull force was required to maintain longitudinal trim in sideslips. This effect was greatest in right sideslips.
Side-force characteristics.- The variation of angle of bank with angle of sideslip, which is a measure of the side-force characteristics, Is presented in figures 15 to 21. The requirement that the direction of bank should always be the same as the direction of sideslip was satisfied in all flight conditions at all speeds.
Effect of sideslip on flap deflection.- The landing flaps on the -F-1 airplane were spring-loaded to permit them to blow up as the
F R
airspeed increased and thus to prevent overloading. Figure 24 presents data obtained during the sideslip tests which show the variation of flap deflection ,,rith angle of sideslip for two power settings at 150 miles per hour.
Sideslip Characteristics with Three Tail Configurations The following discussion will be in the form of a comparison of the directional stability of the airplane with the three vertical-tail configurations described earlier in the report.
Figures 25 and 26 present a comparison of the directional stability characteristics with the three vertical-tail configurations with the belly tank installed. The data of figures 25 and 26 indicate that
either vertical-tail configuration 2 or configuration 3 improved the
directional stability at small sideslip angles where there had been a deficiency with tail configuration 1. In the rated-power, clean condi- tion at 150 miles per hour (fig. 25(a)) where the control-free direc- tional stability in left sideslip$ had been zero with tail configuration 1, there was definite positive control-free stability with either tail con- figuration ? or 3.
Comparison of the curves of rudder force and position against angle of sideslip in figures 25 and 26 shows that tail configurations 2 and produced increases in directional stability. The following table p resents the values of the directional-stability parameters, dSrjdR, the rate of change of rudder position with angle of sideslip, and dFr /dp, the rate of change of rudder control force with angle of side-
slip, measured between 0 0 and 50 with the three vertical tails in two
flight conditions with bell e tank installed at several speeds. The minimum slopes usually occurred between 0 0 and 50 left sideslip.
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NA.CA PM No. L7L31 LATERAL AND DIRECTIONAL CONTROL Rudder to Overcome Adverse Aileron Yaw The ability of the rudder to overcome the yawing moment due to • full aileron deflection with a control-force increment of less than il ^ • : 180 pounds his been discussed in the section on sideslip due to aileron deflection.
Rudder Control in Take-Off and Landing No take-offs or landings were recorded in a 90 0 cruse wind.
Figures 28 to 30 present time histories of take-offs made with tail configurations 1 and 3 and a landing made with tail configuration 1.
No indication of a lack of directional control or of excessive rudder forces was obtained.
Inadeouate directional control in the wave-off condition has been reported. The tests showed that approximately 90 percent of the avail- able rudder deflection was re q uired for trim in level flight near the stalling speed in the wave-off condition with power for level flight with tail configurations 1 or 2. It was found that with tail configuration 3 there was sufficient rudder control to sideslip 20 0 to the left at 85 miles per hour in the wave-off condition at rated power. This amount of sideslip was 10 0 more than that required for trim with wings level and would provide a reserve of control for use in a wave-off.
Directional Control in Straight Flight and in Strafing Runs Figures 31 and 32 present the variation of directional trim charac- teristics over the test speed ranges in the rated-power clean, and the wave-off conditions for all three tail configurations. The change of rudder force with speed between the stall and a speed of 400 miles per hour in the rated-power clean condition for a trim speed of 250 miles per hour was approximately 60 pounds with tail configuration 1 (calcu- lated from the data of fig. 31), 160 pounds with tail configuration 2, 3. The rudder-force change with and 110 pounds with tail confi guration speed satisfied the Navy re q uirement for diving flight. However, a desirable reduction in the force variation with speed could be effected by use of a springy tab.
In order to investigate how well the F8F-1 could hold an aiming point, strafing runs on fixed targets have been made. Two time histories typical of the data obtained are presented in figure 33. Figure 33(a) Illustrates a strafing run which was made by a service pilot in an F8F-1 airplane with belly tank installed and with tail configuration 1.
14 NACA RM No. L7L31 In later tests a gun camera synchronized with the other recording instru- ^•^ ments was used to obtain a record of relative motion between the aiming ^.: point and the target. Figure 33(b) illustrates a strafing run which was made by an NACA pilot in an F8F-1 airplane with belly tank installed and with tail configuration 3. The gun camera and sideslip data of figure 33 show that throughout the range of speeds covered (250 to 400 miles per hoiir) the airplane was subject to pitching and yawing motions of approxi- mately ±0.2 0 amplitude.
Power of Rudder Trimming Tab Data on the variation of rudder-tab angle _°or trim with airspeed are presented in figure 34. A table of minimum directional trim speeds for the various tail configurations follows: NACA RM No. L7L31
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Average I Figure
Power Flaps Landing
I
gear speed
Level
Up 148 36(g)
Up I
flight 36(a)
245 + 36(a)
36(a)
{ 343 135, 36(a)
397 36(a)
Down l Down 87 36(b) I i 136 36 (b) Figure 35 presents time histories of typical aileron rolls obtained at 340 miles per hour in the rated-power, clean condition.
The variation of rolling effectiveness and control force with total aileron deflection are presented in figure 36 for each flight condi- tion and speed investigated. Figure 37 shows the variation of right- aileron spring-tab angle with right-aileiun deflection in aileron rolls with power for level flight in the clean condition at speeds from 148 to 397 miles per hour. The variation with airspeed of rolling velocity at 10,000 feet, total aileron angle, and helix angle pb/27 for a control force of 30 pounds and for full stick deflection is presented in figure 38. The variation of lateral trim force with speed is given
in figure 39
The aileron control characteristics of the F8F-1 airplane may be summarized as follows: (a) The control force required and the rolling velocity obtained In abrupt aileron rolls varied smoothly with aileron deflection throughout the speed range.
(b) The ailerons exhibited no undesirable lag characteristics sad the rolling accelerations were always in the correct directions.
(c) No reversal of rolling velocity due to adverse aileron yaw ever occurred.
(d) Examination of figure 38 indicates that the aileron effective- ness pb/2V in the clean condition with a control force of not over 30 pounds varied from approximately 0.075 at 150 miles per hour to 0.036 at 400 miles per hour. Due to the large reduction of available aileron deflection with increasing airspeed, cne value of pb/2V at 400 miles per hour was approximately 0.046 for full stick deflection. The require- ments of reference 1 with regard to aileron effectiveness were not satisfied.
NACA RM No. L7L31 It would be possible, although difficult mechanically, to increase the aileron effectiveness without changing the control-force gradient by reducing the spring constant and the length of the tab horn in the same ratio. As a result of this modification the available aileron deflection would be greater at all speeds.
(e) The maximum value of pb/2V obtained in the landing condition was approximately 0.08.
(f) There was no tendency for the aileron forces to overbalance.
(g) Figure 39 indicates that the variation of aileron trimming force with speed was small. The aileron trimming tab was found to be adequate.
C O N C L U S I O N S The results of the tests to determine the lateral and directional stability and control characteristics of FSF-1 airplanes (BuAer Nos. 90 1 61 and 94873) with three vertical-tail configurations may be summarized as follows: 1. The control-free lateral and directional oscillations of the airplane were always damped to half amplitude in less than 1 cycle in all test conditions or configurations. However, there was an annoying, persistent, small-amplitude oscillation of the airplane caused by fuel sloshing. Oscillations of the aileron and rudder were satisfactorily damped.
2. The rudder-fixed directional stability was not great enough to restrict satisfactorily the adverse yaw due to abrupt aileron deflection.
The above deficiency was considerably reduced by the enlarged vertical tails of configurations 2 and 3.
3. With tail configuration 1 the rudder-fixed and rudder-free directional stability was low at small angles of sideslip. Addition of the 150-gallon belly tank caused a noticeable reduction in directional stability. The rudder-free directional stability was zero in left side- slips at 150 miles per hour in the rated-power, clean condition with belly tank installed.
The enlarged vertical tails of configurations 2 and 3 increased the directional stability in all flight conditions which were investigated.
However, there was a tendency, toward rudder overbalance at about 150 right sideslip with tail configurations 2 and 3. It is believed that addition of a dorsal fin similar to the one on tail configuration 1 would eliminate the tendency toward overbalance.
NACA RM No. L7L31 4. The effective dihedral, control fixed and control free, was positive except in the power-approach condition at speeds near the stall where the control-free effective dihedral was neutral.
5. The pitching moment due to sideslip was such that a fairly large pull force was required to maintain longitudinal trim in sideslips.
This effect was greatest in right sideslips.
6. The rudder control characteristics were satisfactory for over- coming adverse aileron yaw, for maintaining straight ground paths, and for maintaining straight flight paths with the wings level in all flight conditions at any speed. There was a greater reserve of rudder control available for coordinated turns in the wave-off condition with tail configuration 3 than with the other configurations.
7. The variation of rudder control force for directional trim throughout the test speed range in the rated power, clean condition for 250 miles per hour was about 60 pounds with tail config- a trim speed of uration 1, 160 pounds with tail configuration 2, and 110 pounds with tail configuration 3. Although the rudder-force change with speed satisfied the Navy requirements for diving flight, it would be desirable to reduce this variation of force with speed possibly by use of a springy tab. No excessive rudder forces were required during take-off, landing, or the maneuvers covered in these teats. The variation of aileron force with speed was small. The power of the aileron trimming tab was adequate. The rudder triming tab did not meet the low-speed requirements with any of the three tail configurations but was most adequate on tail configuration 3.
8. There usually were small-amplitude (±0.2 0 ) pitching and yawing oscillations during strafing runs.
9. The response to abrupt aileron deflection was satisfactory, and the aileron rolling effectiveness pb/2P for full stick deflection ranged from approximately 0.08 at q0 miles per hour to approximately 0.04 at 400 miles per hour. The aileron effectiveness was roughly 20 percent below the required value for maneuvering but was satisfactory for control during landing. Full stick deflection could be obtained with a stick force of not over 30 pounds up to a speed of 250 miles per hour (average for left and right roll). Full stick deflection at 400 miles per hour required a control force of 45 pounds.
NACA RM No. L7L31 10. Of the tail configurations tested, configuration 3 was the most satisfactory and would probably be more so if a dorsal fin were aided.
Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics ^^•; Langley Field, Va.
Harold L. Crane
Aeronautical Engineer John P. Reeder Engineer - Test Pilot
Approved-
Melvin N. C..<1ugh
Chief of Flight Research Division bbs R E F E R E N C E 1. .Anon.: Specification for Stability and Control Characteristics of
1945. Fa*u_d74
Airplanes. SR-119A, Bur. Aerc., April 7,
NACA PM No. L7L31
TABLE I GENERAL SPECIFICATIONS OF THE AIRPLANE •••• . Maka and
designation . . . . . . . Gruamman F8F-1 (BuAer Nos. 94873 and 90461)
• Engine . . . . . . . . . . . . Pratt & Whitney R-2800-34-W double Wasp Power ratings: sea level Take-off . . . . . . . . . . . . . 2100 hp at 2800 rpm at Military . . . . . . . . . . . . . 1600 hp at 2800 rpm at
16,000 ft
Normal maximum: blower . . . . . . . . . . . .
Low 1700 hp at 2600 rDm at 7000 ft
High blower . . . . . . . . . . 1450 rpm at 18,500 ft
hp at '?16 " Propeller . Hydraulically controlled four-blade constant-speed Aeroprop Model . . . . . . . . . . . . . . . . . . . . . . . . . A 642 G-1 Blade number . . . . . . . . . . . . . . . . . . . . . . . . .
65o65
Basic Ditch settings, deg. . . . . . . . Maximum 63.0, minimum 28.5
Diameter . . . . . . . . . . . . . . . . . . . . . . . . 12 ft 7 in.
Fuel capacity, gal: .
Main tank . . . . . . . . . . . . . . . . . . . . . . . .
Droppable (belly) . . . . . . . . . . . . . . . . . . . . 100 or 150
D roppable (wings) . . . . . . . . . . . . . . . . . . . . . . . 100
Oil capacity, gal:
One tank (in engine compartment) . . . . . . . . . . . . . . . . 17
War emer gency power system fluid, gal:
One tank (in engine compartment) . . . . . . . . . . . . . . . . 16
General: Span ( wings spread) , ft . . . . . . . . . . . . . . . . . . . .
35.5
Span (wings fol ried), ft . . . . . . . . . . . . . . . . . . . 23.25
Length ( over all) , ft . . . . . . . . . . . . . . . . . . . .
27.5
Length (tail wheel on ground), ft . . . . . . . . . . . . . .
28.25
Height (tail wheel on ground, propeller blade vertical), ft . 13.67
Weight for tents (approx.), lb . . . . . . . . . to 10,000
8,500
NACA NACA RM No. =7 l 21 TABLE I - Continued ^..
GENERAL SPECIFICATIONS OF THE AIRPLANE - Continued .'
'.: Wings:
Area, sq ft . . . . . . . . . . . . . . . . . . . . . . . . . . 244
•
Airfoil section:
Root . . . . . . . . . . . . . . . . . . . . . . Modified 23018
TiD. . . . . . . . . . . . . . . . . . . . . . . . . . . .
23009
Mean aerodynamic chord, in . . . . . . . . . . . . . . . . . . 87.55
Leading edge M.A.C. aft of lead;nq edge of root chord, in.
8.17
R oot chord, in. . . . . . . . . . . . . . . . . . . . . . . . 115.9
Tip chord (6 in. inboard of actual tip), in. . . . . . . . . 51.5
ncidence, deg . . . . . . . . . . . . . . . . . . . . . . . . . 3.0
I
Dihedral, deg . . . . . . . . . . . . . . . . . . . . . . . . . . 5 .5
Sweepback of leading edge, deg . . . . . . . . : . . . . . . . . 5.1
Wing flaps:
otal area, sq ft
T . . . . . . . . . . . . . . . . . . . . . . 18.18
Deflection, maximum down, deg . . . . . . . . . . . . . . . . . . 40
Ailerons:
T otal area, so ft . . . . . . . . . . . . . . . . . . . . . . . 15
Spring-tab area (total), sq ft . . . . . . . . . . . . . . . . . 1.4
Trimming-tab are,e., so ft . . . . . . . . . . . . . . . . . . . . 0.7
Trimming-tab deflection angle, deg . . . . . . . . . . . . . . . t5 Hor'_zontal tail:
Span, in . . . . . . . . . . . . . . . . . . . . . . . . . . . . 189
Total area, sq ft . . . . . . . . . . . . . . . . . . . . . . 52.27
Elevator area (including tabs), sq ft . . . . . . . . . . . . 18.63 Elevator-trimming-tab area (total), sq ft . . . . . . . . . . . 1.0 Elevator-tab'range (approx.), deg . . . . . . . . . . . 8 up 20 down
Tail incidence, deg . . . . . . . . . . . . . . . . . . . . . . 0.5
Vertical tail: Configuration 1 Total urea, sq ft . . . . . . . . . . . . . . . . . . . . .
17.7
Rudder area, sq ft . . . . . . . . . . . . . . . . . . . . . 6.7
Rudder-tab area . . . . . . . . . . . . . . . . . . . . . . . 0.6
Fin offset, deer . . . . . . . . . . . . . . . . . . . . . . . 2.0 Rudder-tab range (apurox .) , deg . . . . . . . . . . . . . . .
f17
Configuration 2
Total area, sq ft . . . . . . . . . . . . . . . . . . . . . 21.7
Rudder area, sq ft . . . . . . . . . . . . . . . . . . . . . 6.7
Rudder-tab area, eq ft . . . . . . . . . . . . . . . . . . . 0.8
Fin offset, deg, . .
. . . . . . . . . . . . . . . . . . 1.5
Rudder-tab range(approx.), deg . . . . . . . . . . . . . . . t17 NACA i
22 NAC A PM No. L7L31
•••• TABLE I - Concluded
•^ G=A.L SPECIFICATIONS OF THE AIRPLANE - Concluded
•• Vertical. tail (concluded):
• Confixirat'on 3
Total area, sq ft . . . . . . . . . . . . . . . . . . . . .
20.7
Rudder area
., sq ft . . . . . . . . . . . . . . . . . . . . . 8.2
Rudder-tab area, sq ft . . . . . . . . . . . . . . . . . . . 0.8
Fin offset, deg . . . . . . . . . . . . . . . . . . . . . . . 1.5
Rudder-tab range (approx.). deg . . . . . . . . . . . . . . . ±17 NACA
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Figure 2.- Concluded.
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NACA RM No. L7L31
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Figure 8.- Oscillations of F8F-1 airplane in the clean condition with power for level flight at 250 miles per hour at an altitude of bOOO 'feet caused by releasing the controls in steady sideslips.
(Tail configuratiQn 3 with belly tank on.)
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Figure 10.- Time histories of lateral oscillations in smooth air due to
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NACA RM No. L7L31
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Figure 10.- Concluded.
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Figure 11.- Time histories of oscillation produced by abrupt left and
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configuration 1; belly tank installed. (2,90 mph)
NACA RM No. L7L31
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configuration 1, without belly tank.
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Figure 15.- Sideslips in the rated-power, clean condition (41 in. Hg at
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closed, F8F-1 airplane with tail configuration 1.
NACA RM No. L7L31
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(b) 245 miles per hour, rudder tab 0.5 0 left, oil cooler closed.
Figure 15.- Continued.
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Figure 1b.- Concluded.
NACA RM ND. L7L31 -r.
(a) 150 miles per hour, rudder tab 14.0 0
left, oil cooler open.
ure 16.- Sideslips in the rated-power, clean condition (41 in. Hg,
-600 rpm) at 5000 feet with belly tank on, cowl closed, canopy
closed, F8F-1 airplane with tail configuration 1.
NACA R11 No. L7L31
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Figure 16. - Continued.
NACA RM No. L7L31
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hour, rudder tab 1.7 0 right, oil cooler closed.
Figure 16.- Concluded.
NACA RM No. L7L31
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NACA 1 1 ! i .
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(a) 150 miles per hour, rudder tab 5.7 0 right.
Figure 17.- Sideslips in the clean, engine-idling condition at 5000 feet
with belly tank off, cowl closed, oil cooler closed, canopy closed, FSF-1 airplane with tail configuration 1.
NACA RM No. L7L31
I I , I { - I i -, I I I
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1 I 1 { I I
I ; I I I I I 1 I 1 I I 1 ^ I I { - I I 3 { NACA - . I ^ i I I (b) 245 miles per hour, rudder tab 5.7 0 right.
Figure 17.- Concluded.
NACA RM No. L7L31
(a) 150 miles per hour, rudder tab 5.7 0 right.
Figure 18.- Sideslips in the clean, engine-idling condition at 5000 feet
with belly tank on, cowl closed, oil cooler closed, canopy closed, F8 F-1 airplane with tail configuration 1.
•••; NACA RM No. L7L31
(b) 245 rr-files per hour, rudder tab 6.3 0 right.
Figure 18.- Concluded.
NACA RM No. L7L31
NACA RM ND. L7L31 (b) 150 miles per hour, rudder tab 0.8 0 right.
Figure 19.- Concluded.
NACA RM No. L7L31
(a) 90 miles per hour, rudder tab 17.b
left (full tab).
Figure 20.- Sideslips in the power-approach condition (20 in. Hg at
2300 rpm) with flaps and landing gear down at b000 feet, with
belly tank on, cowl closed, oil cooler closed, canopy open, F8F-1 airplane, with tail configuration 1.
NACA RM No. L7L31
(b) 150 miles per hour, rudder tab 2.3 0 left.
Figure 20.- Concluded.
NACA RM No. L7L31
..
(a) 105 miles per hour, rudder tab 14.7 0 right.
Figure 21. - Sideslips in the landing condition with engine idling and
flaps and landing gear down at b000 feet, with belly tank on, cowl
closed, oil cooler closed, canopy open, F8F-1 airplane with tail
configuration 1.
NACA RM No. L7L31
•
(b) 150 miles per hour, rudder tab 12.8 0 right.
Figure 21.- Concluded.
NACA RM No. L7L31
Figure 22.- Steady sideslips at 20,000 feet in the rated-power, clean
condition (41 in. Hg at 2600 rpm) at 300 miles per hour with belly
tank on. F8F-1 airplane with tail configuration 1.
a
NACA RM No. L7L31
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Figure 25.- Steady sideslip characteristics in the rated-power, clean
condition (41 in. Hg at 2600 rpm) with three vertical tail con-
figurations. F8 F -1 airplane.
NACA RM No. L7L31
,
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(b) 250 miles per hour.
Figure 25.- Continued.
NACA RM No. L7L31
• I I f i I i I ^ II ^I' I ! I I ^ i i i + ! I ( _ I 1 !
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(c) 350 miles per hour.
Figure 25.- Concluded.
NACA RM No. L7L31
J • (a) 90 miles per hour.
Figure 26.- Steady sideslip characteristics in the power-approach
condition (18-20 in. Hg at 2300 rpm), flaps and landing gear down
with three vertical tail configurations F8F-1 airplane.
NACA RM No. L7L31
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Figure 26.- Concluded.
NACA RM No. L7L31
Figure 27.- Time history of right sideslip at approximately 120 miles
per hour in the power-approach condition (30 in. Hg, 2200 rpm,
flaps and landing gear down) during which rudder overbalance
occurred, F8F-1 airplane with tail configuration three, belly tank
on.
NACA RM No. L7L31
7n, _T 7 v r r v ^ --^3m L- i .
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d
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Figure 28.- Time history of simulated carrier take-off (54 in. Hg,
2800 rpm), by a service pilot. F8F-1 airplane with tail con-
figuration 1.
•••• NACA RI\/T No. L7L31 •• . ..
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t I I _ ! I _L ^_ t I (a) Made by NACA pilot.
Figure 29.- Time histories of take-offs, F8F-1 airplane with tail
configuration 3.
NACA RM No. L7L31
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(b) Made by service pilot. (Elevator and aileron records eliminated
for clarity.)
Figure 29. - Concluded.
....
000000
NACA RM No. L7L31
.
Figure 30.- Time history of simulated carrier landing (22 in. Hg,
2400 rpm), F8F-1 airplane, with tail configuration 1, made by
service pilot.
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NACA RM No. L7L31
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Figure 32.- Directional trim characteristics in the wave-off condition,
flaps and landing gear down and normal rated power (41 in. Hg at
2600 rpm), with two modified vertical tail configurations, F8F-1 air-
plane.
NACA RM N:). L7L31
(a) With tail configuration 1, made by service pilot.
Figure 33.- Time histories of strafing runs, F8F-1 airplane with belly
tank on, oil coolers closed.
• NACA RM ND. L7L31
(b) With tail configuration
3, made by NACA pilot.
Figure 33.- Concluded.
NACA RM No. L7L31
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NACA RM No. L7L31
t .
i • (a) Left roll. (b) Right roll.
Figure 3b.- Time histories of rolls at approximately 340 miles per hour, rated power, flaps and gear up. F8F-1 airplane.
•••• NACA RM No. L7L31
.
h• ..
.. .
(a) Power for level flight, flaps and gear up.
Figure 36.- Aileron characteristics in rolls at 5000 feet, F8 F-1 airplane.
NACA RM No. L7L31
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(b) Power for level flight, flaps and gear down.
Figure 36.- Concluded.
NACA RM No. L7L31
- --- ----- ---- --
Figure 37.- Variation of right aileron spring tab angle with right aileron
angle. Power for level flight, flaps and gear up. F8F-1 airplane.
NACA RM No. L7L31
Figure 38.- Variation of rolling velocity at 10,000 feet, total aileron
angle, and helix angle with indicated airspeed, power for level
flight, flaps and gear up. F8F-1 airplane.
r •• • •
NACA RM No. L7L31
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