Survey
* Your assessment is very important for improving the workof artificial intelligence, which forms the content of this project
* Your assessment is very important for improving the workof artificial intelligence, which forms the content of this project
-%%”-- ~’RM A9101 ,— -g -- .-. _—.. .- .=—-. ‘... - -— .4!s=-4 ~.-. ._ —- .- - NACA “----””’:’-”’--”~ --., Tr,,, ---,--,. -“% “G .-. .-’ *7-RESEARCH INVESTIGATION . l%FOOT ‘J. RANGE OF A THIN WING OF ASPECT PRESSURE STABILITY MEMORANDUM WIND TUNNEL. AN3 CONTROL OF A SEMISPAN By Ben H. Johnson, Ames V - STATIC THROUGHOUT’ MODEL RATIO 4 llff TEE AMES LONGITUDINAL THE SUBSONIC SPEED OF A SUPERSONIC Jr. , and Francis Aeronautical Moffett Field, AIRPLANE W. Rollins Laboratory Caw . ,, -—- .-_:. ~ ---’:1~ ,.----->-—---- . ,“‘, 4 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON December 8, 1949 TECH LIBRARY KAFB, NM NAC!ARM A9101 NATIONAL ADVISORY COMMITI’EXFOR AERomKJTms RESEARCH MEMORANDUM INVESTIGATION OF A ‘IH13Y ~ OF ASPECT RATIo4m Tfm AMEs 12-OOT PRE3SUlZEWII?DTUNNEL. V - STATIC LONGITUINXKG STABILITY AND CONTROL TEROUXOUT THE SUESONIC SPEED RANGE OF A SEMISPAN MODEL OF A SUPEFSONK! AIRX’LANE By Ben E. Johnson, Jr., and Francis W. Rollins STJNMARY A #-. d Wind-tunnel tests have been made of a semispan model of a hype thetioal supersonic airplane to determine the static longitudinal– stabilityand -oontrol Oharaoteristics of the airplane throughout the range of subsonic Mach nmibers up to 0.95. The semispan model had a long slender fuselage and a wing “and horizontal tail of asyect ratio 4 and taper ratio 0.5. The midohord lines of the wing and of the hori– zontal tail were normal to the @ane of symliletry.The profile of the wing and of the tail was a sharp-edged, faired, symmetririaldouble wedge with a thickness-ohord ratio of O.042. Tests were made with the horizontal tail nmunted in the extended wing-chord plane and alternately 6$). 6 percent of the wing mean aero-c chord above the extended wing– chord planet At a constant Reynolds nuuiberof 2,000,000 measurements were made with various stabilizer angles of the lift, drag, and pitching ?mxnentof the model at &ch numbers from 0.20 to 0.95. With the wing * flaps deflected for maximum lift, similar measurements were made at a Mach nuniberof 0.20 with Reynolds nunbers up to 10,000,000. Measurements were made of the dynamic pressure at the two locations of the horizontal tail and of the character and location of the wing wake for the range of l+kchnumbers and Reynolds numibersnoted above. At zero lift~ the Mach nurib er for drag divergence, defined as the Mach number at which the slope of the drag coefficient with respect to Mach number equals 0.10, waa about 0.92 for either looation of the horizontal tail. The angle of attack for a constant lift coefficient decreased slightly with increasing Mach number but no marked or abrupt compressibility effects were evident at lift,coefficients less than 0.6. The contribution of the horizontal tail to the static longitudinal stability at low lift coefficients deoreased with increasing Mach nuder, /’ mH?” 2 RACA RMA9101 primarily due to an increase with Mach number of the rate of change of effective downwash angle with angle of attack. For the model with the horizontal tail in the extended wi&-chord plane, this decrease in the contribution of the horizontal tail to the static longitudinal stability was aggravated by the reduotion with increasing Mach nuniberin the -C-pressure ratio at the tiil. With the horizontal tail mounted in the extended wing-chord plane, static longitudinal stability existed about the quarter point of the wing mean aerodynamic chord at all lift coefficients for &ch nunbers less than 0.87. At ~oh numbers between 0.87 and 0.95, the model was neutrally stable or unstable at lift coefficientsless than 0.30. With the horizontal tail mounted above the extended wing-chord plane, the results indicated static longitudinal s,tibilityat all lift coefficients for all M@h nunibersfor whioh &ta were obtained, For both positions of the tail, either an all+vable stabilizer or a constan~hord elevator provided sufficient longitudinal control to bahnce the airplane at all test Maoh nunibers. t INTRODUCTION . AS a part of a geneml program to determine the subsonio charactez= ii3tics of wing plan forms suitable for flight at supersonic speeds, a series of tests of a thin sha~ged wing having an aspect ratio of k and a taper ratio of 0.5 have been conducted. .The midchord line of the wing was normal to the air stream. Results of these tests have been reported in references 1 through 4. Results of tests at transonic speeds of a wing of identical plan form and similar profile have been reported in reference ~. The purpose of the present report is to summarize the wing data in terms of the static longitudinal+tability and-control characteristics throughout the subsonic speed range of a h~othetical airplane employing this wing. The airplane was represented b$ a semispan model comprising the wing, a slender pointed fuselage, and a horizontal tail geometrically similar to the wing. Force and moment characteristics of the wing, of the wing-fuselage cortibination, and of the complete model with two different tail heights are presented forhiach numbers Up to 0.95 and a Reymolds number of 2,000,000. With the flaps on the wing deflected for maximum lift, simils,rdata are presented for a Mach number of 0.20 and Reynolds numbers up to 10,000,000. The dynamic pressure at the horizontal tail and the location of the wing wake are presented for the wi~-fuselage conibinationfor the same ranges of Reynolds nuziberand Mach nuniber. The tests of the wing-tail-fuselage combinations were conducted with various horizontal+tabilizer settings to investigate the longitudinal oontrol afforded by an all+novable horizontal tail. Data for an identical horizontal tail with a oonstantihord elevator (reference 6) have been used with the wing+usel..agedata to calculate the longitudi~l+ontrol characteristics of the model with a fixed stabilizer and an elevator. -,----. $%‘“-”” NFID*kNii&z A .< t .— m . 3 NACA RMA9101 The effective downwaah angle at the tail, the Mach number at the tail, and the tail efficiency factor are presented herein. COEFFICIENTS AI’lD SYMBOLS The following coefficients are used in this report: % lift coefficient lift () ($s CD drag coefficient * cm pitching+noment coefficient about an axis normal to the plane of symmetry ~ssing through the quarter point of tie wing mean () aerodynamic chord NI T pitching moment qsc t ( ) () Ho= total-pressure-loss coefficient — 9. The followin& s@ols are used in this report: A a 8peed of sound, feet per second b twice the span of the semispanwing, feet c local wing chord, feet ~! wing mean aerodynamic chord, chord through centroid of the ‘; b/2 c2dy f wing semispan plan form b/z H C dy Fo local stagnation pressure in the region of the horizontal tail, pounds per square foot. Ho free+tream stagnation pressure, pounds per square foot it angle of tb stabilizer setting with respect to the wing+hord plane, degrees -Lt tail length, di8tance from quarter point of the wing mean aerodynamic chord to the quarter point of the horizontal– tail mean aerodynamic chord, feet () . . N4CA RM A9101 Mach number (V/a) Mach number at the position c-orrespondingto the centroid of the semitail area normal-acceleration factor of the airplane free-stream dyn&Lo pressure (&), pounds per square foot . dynamic pressure at the position corresponding to the centroid of the semitail area, pounds per sq~re foot Reynolds nuniber pvc ‘ () v area of the semispan wing, square feet area of the horizontal semitail, square feet local airspeed in the.tunnel-floor bou@ary second .— layer, feet per airspeed, feet per second — distance from the plane of symmetry, feet n effective angle of attack of the horizontal tall, degrees J- . . angle of attack of the win&chord plane, degrees tunnel-wall boundary-layer thickness, inches 5 displacement thickness of the boundary layer [f 1 inches (1+/V)dy 1 , elevator deflection, measurd In a plane perpendicular to the elevator hinge axis, positive downward, degrees trailing%dge flap deflection, measured in a plane perpendicular to the flap hinge axis, posi,tivedownward, degrees leading-edge flap [email protected], measured in a plane perpendicular to the flap hinge axis, positive downward, degrees effective average angle of downwash, positive when the air is deflected downward, degrees — . . efficiency of the horizontal tail visco%ity of air, slugs per foot-second . — NKA RMA9101 P mass density of air, slugs per cubic foot 5 MODEL AND Al?PARATUS The tests were conducted in the Ames 12-foot pressure wind tunnel, which is a closed-throat variabltiensity wind tunnel with a lowturbulence level closely approximating that of free air. The steel semispan model wing used for this investigation was the one used in the tests reported in reference 1 and represented a wing of asyect ratio 4 and taper ratio O.~. The midchord line of the wing was perpe~ dicular to the plane gf symmetry. The wing profile was a faired double wedge having a thiclmess+hord ratio of 0.042. The horizontal tail was identical in plan fomand profile to the wing and had an area equal to one quarter of the wing area. Dimensions of the semifusehge and its location with res~ect to the wing are given in figure 1. The semi– fuselage was fitted tightly to the wing and tail without fill.etsat the intersections. For a portion of the tests, tie rear part of the fuselage was modified as shown in figures l(h) and 2(c) to study the effects of such a modification on the yitching+aoment characteristics of the model. The wing was equipped with a full-span, constan+chord, leadingedge plain flap and a 60.9-percen&~n, constantihord, trailing-edge plain flap. The area of the leading-edge flap was 17 percent of the total area of the semispan wing and that of the trailing-edge flq was 12 percent of the total srea of the semispan wing. The unsealed gaps between the flaps and the wing were 0.015 inch with the flaps unreflected, The horizontal tail was mounted in the extended wing+hord plane (figs. l(a) and 2(a) ) and alternately 13 inches (O.6g6c’) above the extended wing-chord plane (figs. l(b) and 2(b) ). To mount the tail above the fuselage, a bracket with a fairing body to enclose the fittings at the point of attachment of the tail surface was added to the fuselage. With the tail mounted in either yosition, provision was made to vary the angle of the stabilizer by pivoting it about its >percen~hord Ilne. As shown in figure 2, the, semis-panmodel was mounted with the wing perpendicular to the floor which served as a reflection plane. me gap between the model and the tunnel “floorwas maintained %etween 0.010 inch and 0.150 inch. No attempt was made to remove the tunnel-floor boundary layer which, at the location of the model, had a displacement thickness 5* of 0.5 inch. The velocity characteristics of the wing+Flselage wake at the longitudinal location of the horizontal tail were measured with a relceconsisting of 61 total~ressure tubes and 3 static-pressure tubes. The rake yas mounted from the tunnel floor with the total-pressure tubes at a position corresponding to the centroid of the semitail area. 6 NACA RMA9101 . CORRECTIONS TO DATA c The data have been corrected for the ef’fectsof tunnel-wall interference, of constriction due to the tunnel walls, and of modelsupport tare forces. The method of reference 7 was used in oomputdng the corrections to the data for tunnel+rall interference. The following corrections were added: &l = 0.363 CL MD = o.oo56~2 -- ACm=O Corrections to the data for the constriction effeots of the tunnel walls have been evaluated by the method of reference 8. The tignitudes of these corrections as applied to Mach number and to dynamic pressure (measured with the tunnel empty) are illustrated by the following table: I Corrected Mach number Uncorrected Mach number I q~orrected quncorrected L Wing alone Wing and Wing alone I Wing and fuselage I fuselage I I 0.95 .92 .90 .85 .80 70 .50 .20 ● 0.937 .915 ; 8&7 “ 799 700 .300 .200 ● ● 0.917 .896 .881 .838 ●7Z .696 kgg 200 . ● I L 005 1.003 1.002 1.002 1.001 1.001 1.001 1.001 L 036 1.027 1.023 1.016 1.OM 1.008 1.005 1.005 The theoretical choking l.kchnuniberfor the wi~-fusel.age conibination was 0.96. Tare corrections due to the air foroes exerted on the turntable were obtained from foroe meam.trementsmade with the model removed from the tunnel. Possible interference effects between the model and the turntable were not evaluated. !lhemagnitude of the measured tare+rag coefficient, based on the wing area, was independent of Mach nuniberand varied with Reynolds nmiber as follows: . -. . NACA RM A9101 7 + Reynolds number CD_re 2,000,000 6,000,000 10,000,000 0.0063 .0057 .0056 The rake of total-pressure tubes and static-pressure tubes used to measure the dynamio pressure at the horizontal tail was calibrated throughout the complete range of Mach numbers, of Reynolds numbers, and of angles of attack of the rake. TESTS Lift, drag, and pitching+noment data have been obtained for the model and its components in the following conibinations: (1) the wing alone; (2) the wing and the fuselage; (3) the wing, the fus”elage,and the tail mounted in the extended wing-chord plane; (4) the wing, the fuselage, and the supporting bracket for mounting the tail above the fuselage; and (5) the wing, the fuselage, and the tail mounted above the fuse-e. At a Reynolds number of 2,000,000 the model was tested at Mach numbers from 0.20 to 0.95. The range of angles of attack for these tests was from -6° to beyond the stall, exce~t at the higher Mach numbbrs where the range was reduced by the limitations of wind-tunnel power and of model strength. At a Mach number of 0.20 the effect of l+ading+dge and trailing-edge flap deflection (bn = 30° and bf = 50°) was investigated at Reynolds numibersof 2,000,000, 6,000,000, and 10,000,OOO. ~S combination of flap deflections was selected upon the basis of reference 2 wherein it waa shown to be the optimum for maximum lift of the wing alone. To determine the longitudinal control which would be provided by an all+novable stabilizer, the model was tested with the angle of the stabilizer varied in2° increments from-lo” to 4° for the m~elwfth the tail mounted in the extended wing+hord plane and from -6° to 4° for the model with the tail mounted above the fusekge. The velocity distribution in the wing-fuselage wake was investigated at a position corresponding longitudinally to the midchord of the horizontal tail (3.508 wing mean-aerodynamic chord behind the quarter point of the wing mean aero@mmic chord) and corresponding laterally to the location of the mean aerot@amic chord of the tail (0.428 wing mean aerodynamic chord from the plane of symmetry). The extent of the survey was sufficient to permit the determination of the dynsmic pressure at either position of the horizontal tail for a range of angle of attack, of Mach number, and of Reynolds numiber. 8 NACA RM A9101 . An index of the figures presenting the results of this investigation is given in the appendix. 7 RESULl13AND DISCUSSION Force and Moment Characteristics The lift, drag, and pitching+mment characteristics of the nmdel and its components are presented in figures 3 through 26. .- Wing alone.- The effects of Reynolds number and of Mach number on the lift, drag, and pitching+noment characteristics of the wing have been reported in reference 1. Data from that reference for a Re~olds nuriber of 2,000,000 at Mach numbers from 0.20 to 0.94 are reproduced herein in figure 3. The data of this figure indicate no large or erratic effects of compressibility up to a Mach number of 0.94. The wing lift-curve slope was 0.062 at a Mach nuuiierof 0.20 and.increased to 0.095 at a Mach nuniberof 0.94. The total movement of the aerodynamic center at zero lift was only about 7 peroeht of the wing mean aerodynamic chord over the test Mach number range. The force and moment characteristics of the wing with various combinations of leading-edge and trailing+dge flap deflections have been reported in reference 2. The data of this reference indicate that a leading+dge flap deflection of 30° and a trailing+dge flay deflection of 50° were optimum for maximum lift. Data obtained with this combination of flap deflections are presented herein in figure 4 for a Mach numiberof 0.20 and Reynolds numbers from 3,000,000 to 10,000,000. These data show that deflection of the flaps increased the msximum lift of t@e wing from 0.76 to 1.4o and that the aerodynamic characteristics of the wing with the flaps deflected were little affected by increase of Reynolds number to ‘ 10,000,000. c The variation with angle of attack of the lift coefficient of the wing with the gaps sealed and faired is presented in figure 5 for a Reynolds number of 1,000,000 for &ch numbers up to 0.94. Since the wing and tail were geometrically similar and the mean aerodynamic chord of the tail was one-half that of the wing, these data may be.considered to represent the lift characteristics“ofthe isolated tall and may be applied as the characte~ istics of the tail on the model at a Reynolds nuuiberof 2,000,000, based on the wing mean aerodynamic chord, if corrections are made for the downwash and reduction in the dynamic pressure”at the tail. The force and moment cha~cteristics of Wing-fuselage .conibination.– the wing-fuselage combination with the flaps neutral are shown in figures 6, 7, and 8. Comparison of these data with those of figure 3 reveals that addition of the fuselage caused an increase in the dragj a reduction in. .- , .: ...= RACA RMA9101 9 the maximum l.if t at Mach numbers less than 0.80, and a forward movement of the aerodynamic center at low lift coefficients. The lift, drag, and pitching+noment characteristics of the wing-fuselage combination with the wing flap deflected are presented in figure 9. Comparison of these data with those of figure 4 indicates that the addition of the fuselage caused a decrease in the maximuu lift coefficient from 1.40 to 1.34 and an increase of 1° in tie angle of attack for zero lift. The characteristics of the winf-fusehge combination were little affected by a change in Reynolds number from 6,000,000 to 10,000,000, but .anincrease from 2,000,000 to 6,cKK),000resulted in a sizable decrease in the *g. wing, fuselage, and horizontal tail in the extended wing-chord plane.Liftj drag, and pitching+noment characteristics of the cmnplete semispan model with the horizontal tail mounted in the extended wing-chord plane are presented in figures 10, 11, and 12 for Mach numbers up to 0.95 and stabilizer angle settings from 4° to —10~. At a Mach number of 0.20, the aerodynamic center was shifted from 14 percent to 41 percent of the wing mean aerodynamic chord due to the addition of the tail. (See fig. 12(a).) As the Mach nurriber was inoreased, the stabilizing effect of the horizontal tail was diminished to the extent that at a Mach nm?iberof 0.95 the horizontal tail made little or no contribution to the stability of the model at lift coefficients between LO.3. As will be discussed later, this decrease in the contribution of the tail to the stability was due to an increase in &/&L and to a decrease in the dynamic-pressure ratio at the tail as the &ch number was increased. With a stabilizer amglesetting of 0° and in a range of lift coefficients of about kO.30, the complete model was neutrally stable about the quarter point of the wing mean aerodynamic chord at a Mach nuniberof about 0.87 and longitudinally unstable at higher &ch numbers. At lift coefficients greater than 0.30 stability existed at all test Mach numbers. The all-movable stabilizer provided sufficient longitudinal control to balance the airplane model at all Mach numbers up to 0.95 and at all angles of attack up to the stall. The value of (%/~it)CL.O was approximately ~.036 at a Mach number of 0.20 and increased slightly with increasing Mach number. (See fig. X2. ) The lift, -g, and pitching+.oment characteristics of the complete semispan model with the wing flaps deflected are presented in figures 13, 14, and 15 for a Mach number of 0.20 and Reynolds numbers of 2,000,000, 3,CO0,000, and 10,030,000. At lift coefficients from zero to the maximum the complete model was longitudinally stable about the quarter point of the mean aerodynamic chord. Wing, fuselage, and horizontal tail above the extended wing-chord J2sQs” – To investigate the improvement in longitudinal stability and control afforded by raising the horizontal tail above the wing wake, tests were conducted with tie-model tail mounted 13 inches (O. 696 WinF-mean aerod~amic chord) above the extended wing+hord plane. 10 NACA FM A9101 . Mounting the tail above the fuselage necessitated a supporting bracket with a streamlined body to serve as a fairing for the fittings by which the stabilizer was attached. The force and moment cheracteristios of the wing and fuselage with the bracket and the fairing body are presented in figures 16, 17, 18, and 19. These data indicate no noticeable effects of the bracket on the characteristics of the wing-fuselage combination ‘ except a slight increase in the minimum hag. (See fig. 17.) Lift, drag, and pitching+oment characteristics of’the complete semlspan model with the horizontal tail mounted above the extended wingchord pkne are presented in figures 20, 21, and 22 for Mach numbers up to 0.95 and for stabilizer settings from 4° to -@. Comparison of the drag data of figure 21 with those of figure 11 indicates a slight increase in the minimum dmg which maybe attributed to the addition of the tail bracket and the fairing body and not to the raising of the horizontal tail. The model with the high tail was longitudinally stable at all lift ooefflcients below the stall and at all Mach numbers, as can be seen from figure 22. At a Mach number of 0.20, addition of the horizontal tail shifted the aerodynamic center from 14 percent to 53 percent of the wing mean aerodynamic chord. The contribution of the horizontal tail to the longitudinal stability decreased with increasing Mach nuniber. As will be discussed later, this reduction in the contribution of the tail to the st&bility was due -primarilyto an increase in &/A with increasing ldach number. The ~ll+novable stabilizer retained effectiveness in longitudinal control at all Mach nunibersand all lift coefficients. d — F There was,a marked change in the pitching+oment coefficient at zero lift as a result of raising the tail above the fuselage. Whereas with the tail In the extended wing+hord plane, zero pitching moment occurred at zero lift with a stabilizer angle of 00, with the.tail raised above the extended wing-chord plane a stabilizer setting of approxktely 2° was required to produce zero pitching moment at zero lift. To investigate the , cause of this shift in the zerc+lift pitching+noment coefficient the Reynolds nuniberwas Increased from 2,000,000 to 12,000,000 while the Mach ntier remained 0.20. !Thisincrease had no effect on the pitching+ncment coefficient at zero lift. Visual observation, by means of tufts, of the flow at the afterend of the fuselage and on the tail-supporting bracket revealed a sizable stresm angle in the region of the tail due to the rapid convergence of the rear end of the fuselage. This convergence was reduced bymodifylng the afterpart of the fuselage as shown in figure l(b). The results of tests with the modified fuselage are shown in figure 23. These data show that, for the model with the tail mounted above the extended wing-chord plane, modification of the fuselage caused a decrease in the zerc+lift pitching+mne nt coefficient greater than the Increase accompanying the raising of the tail on the otiginal fuselage. The lift, drag, and pitching+noment characteristics of the complete semfspan model with the high tail and the original fusehge and with the wing flaps deflected are presented in figures 24, 25, and 26. Raising . . — . 11 NACA RMA9101 the tail a%ove the fuselage had little effect on the lift and drag of tie model with the flays deflected. However, the model with the high tail had more nearly linear pitching+mment characteristics than the model with the tail in the extended wing-ohord plane. Wing Wake and Effective Downwash at the Horizontal Tail The dymunic pressure at we horizontal tail, tie velocitytistiibution in the wake of the wing-~ elage cabination, the effecti~e angles ~of’d~wash at the horizontal tail, and the tail efficiency factors are presented in figures 27 through 36. Looation of the wiw wake .- The location of the point of maximum totalpressure loss and the wake boundaries have been determined from measurements of the stagnation pressure behind the wiu-f uselage combination at a position corresponding longitudinally to the midchord of the horizontal tail (3.508 wing mean aero@namic chords behind the quarter point of the wing mean aerodynamic chord) and laterally to the mean aerodynamic chord of the horizontal tall semispan (O.128 wing mean aerodynamic chord from the ylane of symmetry)s The results of these measurements are Tresented in figures 27 and 28 where the location of the wake is presented as a function of angle of attack for various Mach numbers and Reynolds numbers. The location of the wake is given with respect to the wing-chord plane at 0° angle of attack. The two alternate positions of the horizontal tail are also identified in these figures so that the location of the tail with respect to the wing-fuselage wake can be readily detemnined. The tail mounted in the attended wing-chord @ane was in the wake of the wing at all test angles of attack and at all test Mach numbers. The high tail did not enter @e wake until the angle of attack exceeded about 70 at l&ch nuuibersbelow 0.70. As the Mach number was increased above 0.70, the high tail entered the wake at progressively lower angles of attack. With the wing flaps deflected the high tail was above the wake at all angles of attack. (See fig. 28. ) . At moderate to large angles of attack and at Mach numbers above 0.85, the wing-fuselage wake was characterized by two distinct regions of hrge total-pressure loss. These are shown in figure 29 which presents the variation of total-pressure loss across the wake at an angle of attack of GO and a ~ch n~er of 0.85. The secondary peak of total-pressure loss is believed to be associated with separation at the ting leading edge and usually occurred near the amgle of attack at which the aerodynamic center of the wing moved forward. Figure 29 also indicates ‘chatthe presence of the fuselage influenced the msgnitude and the location of the total-pressure losses and the location of the wake boundaries. . 12 NACA ~ ,. .. . ....- A9101 . Dynamic-press~e ratio and Mach number at the tail.- To determine the ratio of the dynamic pressure at the tail to the free-stream dynamic pressure, measurements were made of the stagnation and static pressures in.the region of the horizontal tail, The results of.these measurements are presented in figure 30 for various free-stream Mach numbers as a function of angle Qf attack. The dynamic-pressure ratia at the centroid position of.the horizontal tail inthe extend6d wing-chord plane for 0° angle of attack varied from 0.945 at a“free+tream Mach number of 0.20 to 0.865 at a free+t+ream Mach number of 0.95. Due to the symmetry of the model about the wing+hord plane, the dynamic-pressure ratio at the tail_ mounted in the extended wing+hord plane increased with increasing or decreasing angle of.attack, attaining a value of approximately 0.98 at all Mach nunbers at angles of attack of L60. r At a Mach number of 0.95, the dynamic-pressure ratio at the centroid position of the high horizontal tail was unity at angles of attack less than 2.5° and less than unity at larger angles of attack, (See fig. 30(b).) As free+tream Mach number decreased, the minimum angle of attack for which the dynsmic pressure remained at the free-stream value increased to ~ for Mach numbers less than O.~0. With the wing flaps deflected, the dynamic-p?essure ratio at the high tail position was unity, and at the position of we tail in the extended wing-chord plane it varied from approximately 0.99 at 0° angle of attack to approximately 0.84 ~t 10° angle of attack. me effect of increasing the Reynolds number fra 2,000,000 to 10,000,000 was to increase the dynamic-pressure ratio,approximately 5.5 peroent at an angle of attack of 10o with less effect as the angle of attack was reduced. . . — ,_= 6 ~s The Mach numbers at the tail have been computed from the wakesurvey data and are presented as functions of angle of attack for various free+tream Mach numbers in figure 32. Effective angles of dcwnwas’hat the tall.- The effective angles of downwash at the horizontal tail have been computed from the moment data and are presented as avemge values over the stabilizer angle range in figures 33 and 34. The e@ression used for calculation of the effective angle of downwash is as follows: E = where (~q )a u+‘t - Pf%%)= — (*Jai.& is the increment in pitching+uoment coefficient due to the addition of the tail for a constant angle of attack and (~~it)a is the stabilizer effectiveness at a constant angle of attack. This expression does not permit the downwash due to the wing to be sepamted from the downwash due to other components of the model, and thus the stream angle at the horizontal tail due”to convergence of the rear end — . .._ — 13 NACA BMA9101 of the fuselage is Included in the the data. value of the downwash computed from Efficiency of the horizontal tail.- The tail efficiency factor computed from the force and moment data is presented in figures n(@) 35 and 36. The tail efficiency factor, defined as the ratio of the lift produced by the tail in tie presence of the fuselage to the Mft poduced by the isolated tail operating at the same Mach nuniber,was oonqmted by means of the following expression: where (d~/ti)t is the lift-curve slope of the isolated horizontal tail operating at the free-stream Mach nuder of the horizontal tail (figs. 5 and 32). No attempt was made to separate the effects of dynsmic~ressure ratio at the tail from the tail efficiency due to the possible large The tail efficiency factor is variation of q+/q alon8 tie tail s~n. -u. presented as a function Ef Mach number in figure 35. For either position of the tail with the flaw neutral, the tail efficiency factor was less than 80 percent and varied - approximately 10 percent over the test range of Mch nunbers and angles of attack. . The Effects of Compressibility -d The effects of compressibility on the lift, dreg, pitching moment, in figures 37 through and downwash of the complete model are s~rized 46. Lift and tiaq.- !Fhevariati~n with Mach number of the angle of attack for a constant lift coefficient tis -11 (fig. 37), increasing Mach nuniberusually being accompanied by a decrease in the angle of attack for a given lift coefficieq?t. The variation with -h nuniberof the drag coefficient for several constant lift coefficients is shown in figure 38. At a lift coefficient of zero, the bag coefficient of the model with the tail in the extended wing+hord plane started to inorease at a Mach number of about 0.80. For the model with the high tail, the drag increase started at a Mach number of about 0.75. The Mach nuniberfor drag divergence, defined as the l.kch number at which (~D/&)%a = 0.10, was approximately 0.92 for the mbdel with either tail position. Static longitudinal stability and control.- The variation with~ch number of the pitching-ment coefficient for several constant lift coefficients is showr-in figure 39. h general, the pitching+umnent coefficient inoreased with increasing Wch nuniber. The static longitudinal . . NACA RM A9101 ● instability at Mach numbers above about 0.85 of the model with the tail in the extended win&chord plane, as mentioned previously, is evident from the data of figure 39(a). v The variation with Mach nuniberof the effeotive angle of downwash at several oonstant values of the lift coefficient is shown in figure M, and tie variation of ~@x wit@ Mach nuniberis shown in figure 41. For either location of the horizontal tail, &/& increased with .increasing Mach mmiberbut the value of ac/~ and the rate of increase with Mach numher was much larger for the model with the tail in the extended wimgchord plane. The static longitudinal instability at high subsonic Mach nuniberswith the tail in the extended wingahord plane was principally a result of this large value of a+h. The variation with Mach number of the lift coefficient for balance about the quarter point of the wing mean aerodynamic chord is presented in figure 42 for various angles of stabilizer setting. The model with the tail in the extended wing-chord plane was neutrally stable at a Mach nurber of 0.86 and unstable at higher Mach nuniberswhen the stabilizer setting was Oo. With a stabilizer setting of -1° or -o the model was longitudinally stable, but the lift coefficient for balance varied erratically with Mach nuniberat Mach numbers above about O.~0. With the tail mounted above the extended wing-chord plane, the model possessed static longitudinal stability at all stabilizer settings and all Mach nunibers. For positive values of lift coefficient, the balanced lift coefficient for a given eJtdxLlizerangle increased as the Mach number was increased to about 0.90 and decreased with further increase in the Mach nuniber. .L. - For the model with either position of the horizontal tail, the allmovable stabilizer required only 4° to 60 of deflection to balance the nmdel at the stall with the flaps up. The experimental results of this investigation have been used to ‘ predict tie static longitudinal~tability and-control characteristics of a @pthetical airplane with a wing loading of 100 pounds per square foot in flight at an altitude of 10,000 feet. The airplane center of gravity has been assumed to be on an axis perpendicular to the plane of symmet~ pass,ingthrough the quarter poititof the wing mean aerodynamic chord. The variation of airplane lift coefficient with Mach nuder for several values of normal-acceleration factor is presented in figure 43. The calculated effects of fligh~th curvature on the flow at the tail were negligible for the assuniedflight condition. The variation with Mach nuniberof the stabilizer angle required to balance the airplane is shown in f@ure 44 for severialvalues of nor’malacceleration factor. With the horizontal tail in the extended wi~hord plane, the airplane would be longitudinally unstable with a normal- — 15 acceleration faotor of unity at Mach numbers above about 0.87. Below this Maoh nuniber,the variation of stabilizer angle tith speed was stable and a total change of stabilizer angle of 1.P would be neoessary to balance the airplane in level flight between Mach nunbers of O.n and 0.87. With the tail mounted above the extended.wing-ohord plane, the airplane would possess statio longitudinal stability at all Mach numbers but the variation of stabilizer angle with velooity would be unstable at &oh nuxibersabove about 0.90. A change of 2.4° in the stabilizer anglewould be requiredto balancethe airplanein levelfli@t between Mach nuuibers of 0.50 and 0.95. To compare the longitudinal mntrol afforded by the all+novable stabilizer with that whfoh oould be accomplished with a fixed stabilfzer and an elevator, elevato~f femotivenessdata from referenoe 6 were applied to the hypothetical airplane. The tail model of reference 6 was equipped with a 2C@percent area, constan&ohord elevator and the plan form and profile were identical with those of the horizontal tail investigated herein. me elevato~ff activeness data of reference 6 are reproduced herein in figure 45 and in application of “the data it was assumed that there was no effect of scale between Reynolds nmibers of 2,000,000 and 1,000,000 and that the elevator efficiency factor was 100 percent. The variation with Mach number of the elevator deflection required to balance the airplane at the previously assumed flight conditions is presented in figure 46. The calculated static lon@tudinal stability and control of the airphne with a fixed stabilizer and an elevator are similar to those previously discussed for the airplane with the all+aovable stabilizer. About 5&percent greater deflection would be required of the elevator to produce the same balanoe lift coefficient as the all+ovable stabilizer. Longitudinal Characteristics with the Flaps Defleoted The variation with ‘Mft coefficient of the stabil.lzerangle required to balance the model with the flaps deflected is presented in figure 47 for the model with the horizontal tail in the extended wing-ohord plane. The corresponding &’ag ooeffioient is shown in the same figure and the liftArag ratio as a function of lift ooefficien~ for balance is shown in figure 48. . These e~erimental results have been used to preallottie powe=f gliding speed and sinking speed at sea level of a hypothetical airplane wiiiha wing loading of 100 pounds per square foot. The effects of the proximity of the ground and the increased drag due to landing gear have 16 NACA RM A9101 . been neglected. The results of these calculations are presented in figure 49. The minimum powe~ff sinking speed was k6 feet per second and occurred at a forward speed of In miles per hour. SDMMARY r OF RESULTS The results of wfnd-tunnel tests at Mach nunibersup to 0.95 of a semispan model of a ~othetical supersonic airplane with the horizontal tall mounted alternately in the extended wi~hord plaqe and 0.6g6 of the wing mean aerodynamic chord above the extended wing+hord plane have been presented. A summary of these results follows: . 1. At a lift coefficientof zero, the Mach number for drag divergence was about 0.92. There was a smooth increase of lift+.cmrve slope with increasing Mach number up to a Maoh nuniberof 0.95. 2. The contribution of the horizontal tail to the statio longitudinal stability decreased with increasing Mach nuniber. lMis decrease was due primarily to the increase with increasing Mach numiberin the rate of change with angle of attack of the effective angle of downwash at the tail. With the horizontal tail in the extended wh&chord plane, a further destabilizing effect was the decrease in dynamic-pressure ratio at the tail with increasing Mach number. 3. With the horizontal tail in the extended wing-chord plane, the model was longitudinally unstable at Mach nunibers above 0.87 at lift coefficients less *n 0.3. With the horizontal tail 0.696 of the wing mean aerodynamic choziiabove the extended whg-herd plane, the model was longitudinally stable at all lift coefficients for all.Mach numbers for which data were obtained. 4. Eitier anall~vable stabi~zer or a f~ed stabilize rwi~ a constan-herd elevator provided sufficient longitudinal control to balance the model throughout the test range of &ch numbers. Ames Aeronautical Labomtory, National Advisory Committee for Aeronautics, Moffett Field, Calif. . , ‘ 17 NACA RM A9101 . APPmimX The follmd.ngtableshave been includedto providea convenientink K to the figures presentingthe resultsof this investigation: FORCEAED MQ4RTT ~TICS wing Results presented a,@, &Cmvz Flap defl.kction C~ 00 a,~,&hVSCL Alone Reynolds Mach number 0.20 to 2X106 3 3xlo*to lo%lo= 4 0.94 0.20 on=30°,*-500 Figurenumber number W- Alone With All.Gaps Sealed Results presented Flap deflection --- CL Ys a Mach number 0.20 I to 0.94 Reynolds number Figurenumber Moe 5 W~Fuaelage Cabination . . * FIEP I deflmtinn Remits presented I 00 CLvsa 0.20 1 CLVSQ CL W CHL on=300,&=500 a,q)&cmvscL Results presented Mach number map deflectlml to 0.95 I Reynolds number I Figurenuniber 6 2X1CP 7 I i 8 0.20 2xlo= to lmlC@ 9 Stabillz= angle number 40 to -MO 0.20 Figurenumb= cLvsa CL PE ~ CLVZ~ CL~a CL VE 5n=30°,%-500 2X1(Y to lmlo= ~ cLvsc~ 1 i I I is(a) to 13(C) 14(8) to M(c) ~(a)tou(c) NACA RMA9101 WiwFuEelage Combination with Bracket for Mounting Tail Above Fuselage r FlaP deflection Results presented Reynolds Mach number 0.20to 0.95 00 Figure nmber number I > — 2X1O= I 16 17 18 I Woe 0.20 Ei&300,5f=500 Wing, Fuaelag%and Horizontal Tail Above Extended Wi@hord Re6ult’a preaepted CLva F1.q deflection a 00 Mach number Stabilizer angle 40 to-60 Reynolds number 2)(106 2.20to 0.9! 19 KDaoe to Plane Figure number h 20(4 4 to 20(h) CL V8 CD 21(8) to 21(h) CL VS Cm 22(a) to22(h) %Lva 00 cm cL Vs a 0.20, 0.92, 40,00,&-W’ &=300,6f=500 23 acloe 0.90 0.93 ZRIIYto 10X1OC 0.20 24 cL va CD 2,5 CL V6 cm 26 Slmwathe effectof modifyingthe rearof the fuselage. FLOW CONDITIONS IN THE REGION OF TEE ECRI,ZONTALTAIL Chmacteriatica of Win&?uaekge Results presented Location of va a wake FlaP deflection o“ Mach number Wake Reynolds number 0.20to 0.95 2KL06 Figure number 27(E4t027(g) I Z$l=300,5f==oo Pressure loaa In wake vs distance from win~hord plane 00 0.20 0.85 2X1O= to loxlo~ 2X106 28 29 I . 19 NACA RM A9101 ~LT d R-=tio, Meeh Shmhr, ~=e m deflection aeto 0.20 the to 0.9s mil Hgum mmber 30(a) & 30(b) 31 llxuoe 2%UP 32(w)& 32(b) 0.20 to O.* 2xLOe 33(a)& 33(b) 0.20 2x10’3to lmloe 34(E)& Sk(b) 0.20 to o.g5 2XKP 35(.) & 35(b) 00 0.20 00 8U=@,8#30 00 at we toO.* 0.20 8n-3@,8# DOlumaeh ReymJI& number HELoh 00 of 2xfe0tiv0 Angb 0.20 %=3dhf-60° 36(a) & 36(b) 2uo=tolo%loe sLM4#rtY CORVE!3 Zffeots of Cmuressibility cm the Chareoterietfcsof the Mcdel de~lectfon,d; Reynoldn Iu&ber, W] The [KL8p Lift emidrq { Reeults Idft Coefficient me8ented ava~ O to ~TaM Stubillmr engle Maoh number 0.6 J me numbcm 0.20to o.g5 37(a)& 37(b) i 38(E)& 38(b) r stnbtli~ EuMiIxmkYl ehuaateristical: Lai8itudinel .* Resultm Stabilizer I&t caefficklt prenmted .’ o Meoh number 0 0.6 3g(a) b(a) ~ 41 T 00 to -r --- I --- [ 0.50to -----I ----- --- I it for ~.~ qp3 Klap deflection & 0.20 --- Mlgitudinal ChareotertitloaWIReeults preeentea Mach number requir-tm ofby@aletiodeirplme With au- d 0.20b O.gk ks 0.20 h’6 to o.g5 nmbac a b7 wing UaMng “1 : of103pmvMs perOquere footin lo,wJ feet. npeedfor bypothetiod airp@ae vith flight at neaI.mel.Poueroff. %Mnklng 4.4 lctdo= Kpeed m ●med ~ o.g5 R=- 0.2U ‘-’111! at 43 ,q+o 7 zdft -t to k?(a) & &(b) O.* the r-ley Defleotea L/l)Y# CL for %nklng miung ~g(b) k b(b) & a wing I.oeMng of loopomde~B-foOt”in 20 NACA RM A9101 .. REFERENCES 1. Johnson, Ben H., Jr.: Investigation of a Thin Wing of Aspect Ratio 4 in the Ames 12~oot Pressure Wind Tunnel. I - Characteristics of a Plain Wing. I?ACARMA8D07, 1948. 2. Johnson, BenH., Jr., and Bandettini, Angelo: Investigation of a Thin Wing of Aspect Ratio 4 in the Amss lE%Foot Pressure Wind Tunnel. II -The Effect of Constant-Chord Leadi~ and Traili~ Edge Flaps on the Low~peed Characteristics of the Wing. NACA RMA8F15, 1948. 3* Johnson, BenH., Jr., and Demele, Fred A.: Investigation of a Thin Wing of Aspect Ratio 4 in the Ames l~oot Pressure Wind Tunnel. 111- The Effectiveness of a Constan~hord Aileron. NACA RM A8117, @l.8. 4. Johnson, Ben H., Jr., and Reed, Verlin D.: hvestigatfon of a Thin Wing of Aspect Ratio 4 in%he -s l~oot EtressureWind !lhmnel. rv– The Effect of a Constant<hord Ieadin@!ldge Flap at High Subsonic Speeds. NAcARMA8K19, 1949. 5* Rathert, George A., Hanson, Cal M., and Rolls, L. Stewart: Investigation of a Thin Straight Wing of Aspect Ratio 4bythe NhCAW~ F1OW Method. -Lift and Pltchin@40ment Characteristics of the Wing 1949. Alone. NACARMA8L20, 6. Bandettini, Angelo, and Reed, Verlin D.: The Aerodynamic Characteristics Throughout the Subsonic Speed Range of a Thin, Sharp+l!klged Horizontal Tall of Aspect Ratfo h Equipped with a Constant-Chord Elevator. NACARMA9E05, 1949. Sivells, James C., and Deters, Owen J.: Jet~oundary and PMorm Corrections for Partial-Span Models with Reflection Plane, End Plate, or No End Plate in a Closed Circular Wind Tunnel. NACA Rep. 843, 1946. Herriot, JohnG.: Blockage Corrections for ThreeAlime~ional-low Closed-Throat Wind Tunnels, with Consideration of the Effect of Compressibility. ItACARMA7B28, 1947. . k — Af/ dimensions given h jnches unless otherwim specifiti. .—. — . +. ffote:hwdi~edge radii are i2 and trai7hg0.005 inch ;* —-~-— Sym ~-Wing and tail section ‘! &nvdiJ (Fan=#lewi#, Fus.abLw 36.00 (u) Figure l.- Semi- horizontal modsi fal/mwmfedin of on iw~bns extended wing-chord with a wtng and all- plane, movabk horizontal tail of a~cf ratb 4. I-i) P NACA RM A9101 22 All dimensions given in inches unless otherwise specified ., horizontal toil above the fuselage ~ of horizontal Original fusehge, -------- .//, R toil Modified fuseluge 7 I \ ------ -7-:=’--+:.. , -1 ●. . ._. Modified fuse/oge coordinates perce~~ length of “{ originu/ fuselage ) 7 L Toil foiring body coordinates [percent length) r, x, 7.% /4.8/ 22.22 29.63 37.04 50.00 62,96 70.37 ?7.78 95./9 92.59 0000 4.$6 /4. 58 /8, 48 2Q 78 21.85 22.22 2/.85 20.78 /8.48 /4. 58 4.96 (b)Hori;ontol /. - Concluded. r= 0 toil mounted moolfieo’ fuseloge. Figure x= 425 6 5 i97: 7:9; 3.960 8220 8$?09 3.890 3.822 9590 /02.80 2897 2.040 /0825 10S?60 /.746 ///.00 /.438 112.30 /./03 //3.70 0.758 above the original un~ the 1“ , (a) Figure 2.- Semlsp+m model Horizontal tall nmnted of the airplane nmnted In the extended 3n the AJIBS M-foot wing-chord pressure plane. witi tmmd. .—.. .-. . .—. —.— # ... ,., (b) F@ur(j 2.- Conti.wd. Eorizontil ,,, tail mmnted ,, Fdd1::, above the I%ebg.. ~k-. .- .-d . . I“ i i;’ ‘, (C) mgure 2.- continued. Modified fueelage with the horizontal tail munted above the ~elage. . . . E! (d) F@ure 2.- Conoluded. Ralw of premsura tubes muntedkmhirtt hew~uselage conbinatl.on at the location of tho horizontal tail. . . . -. , ● r b ii! Ifvldl I I I I I v b~ d -,4 -.6 E 1- .fjj_ 04 ~ Figure 3.- D8 ./2 for”iW, .16 coeftkien~ 0.20 -4 0 4 8 Angle of offock,a, CD 71e lift, drag,ond -8 pdching-nwment churocteriktics /2 for M, deg of the ploin 0.20 ./ o + PitcMg-moment wing. /?, 2,000,000. forluo20 coa%bieti, - ‘ Cm . c+ . s * $ \ 9) o Q Q .. -1 0 .04 a08 Dtvg ./2 .16 coet%kw~ .20 .24 CD .28 .32 -8 .36 .40 of otfmk, -4 Angh Figure 4.- Tbe lift, drag, and pitching-moment M, 0.20. choructeristics 0 .44 4 8 12 16 a, deg of the plain 20 J wing with fh o + -2 RMJkwm@ flops deflected -IW &, 30°; “,3 L +, 50-; g H o P * , Q. I ,, ,, . 1 NACARM A9101 33 /.2 M /.0 O 0.29 ❑ 0.50 0.70 0.80 0.85 0.90 0.9/ 0.94 0 A .8 V v ~ .6 . v } c -.4 n h d -.6 ‘.8 -/. o . -1.2 -/6 -/2 -8 -4 0 Angle of Figure 5.- The lift characteristics attack, 4 8 a, /2 for A4=0.29 deg of the plain wing, gaps sealed. R, ~000,000. 34 NACA RM A9101 . ● . ‘--8 Figure 6,– -4 0 Angle of uttuc~ The lift chorucferistlcs 4 8 /2 /6 for M=O.20 a, deg of the wing and fuselage. /?l 2,000,000. . NACA RM A9101 35 ...... . . . . — Figure Z— The dreg characterisflcs of the wing und fuselage. R, 2,00~000. NACARM 36 A9101 . /.2 M o /, . 0 0.20 13 0.50 00.70 ~ 0.80 v 0.86 B 0,90 Q 0.93 F 0.95 .8 .6 .4 us’ “ 2 . v -.4 -.6 ,/ -,/ o Pitching-moment Figure 8+ fuselage, The pitching R, 2,000,000. for coefficient’, -moment M=O.20 T Cm chorocterlstlcs of the wing and . , E! I!! mzqivehwm,~ -8 -4 An@ 0 of 4 ott@k, 8 @ 16 20 a, d~ 71 Ru Figww 9,- +, 50”; The Iift,drag,and M,o.20. pitching-moment choractetisiics of the wing and fusebge wi?hthe flaps .2. WemYk@ deflected f%! an, .30”; .... -Q/# 38 NACARM A9101 . . v -4 * 1/ . ! WwFFFFm 1 , ! m m I I I .&? -8 -4 0 &g/e 4 1 I I I { /2 8 16 t I 20 I for I I & = 4° Of 0ttc7Ck, a,deg (U)M, o. Za Figure /0,–rhe lift chart?ctert’sties mounted in f~e extended wing-chord of the airplune plane. model R, 2,000,000. with the horizontal foil 1 39 NACA RM A9101 /.0 .8 .6 .4 H H 00 A ‘2 ‘-ol%ir D -G 4 -8 v -/0 LHii5EE u-LK -.4 L v JT 1/ YY -.6 [ I 1 r I 1 1 1 1 P 1 t -.8 -8 0 Angle of attack, -4 48 u,deg /2 (b) M, O.50. Figure /O.- Continued. /6 for ~= 4“ TYH 40 NACARM A9101 . . * N* .? Q $ Q u $ 4 -8 -4 4 Angle of attack, 0 8 a,09g /2 /6 for 4 “4” (c) M, 0.70. Figure /O.- Continued. . .- NACARM A9101 41 . . /.2 /.0 .8 .6 . . -.2 1/ -.4 1A 141 I I t I I I I I .= -8 -4 Angle of 0 affack, 4 8 /2 a,o’eg ld) Figure 10.– GOnI%7U@d. 16 M, 0.80. ‘ 20 for ~ “ 4° I i I . 42 NACA RM A9101 . . /.2 +- /.o 8 .6 A ‘2 l-i v -4 II v 4I -/0 ,4 * /1 /1 v F‘ / iv P ? .“ -8 -4 0 Angle of attack, I 8 I -=zi=- /2 /6 for ~= 4“ u,deg (e) M, 0.85 Figure /0.– Continueo’I . . NACA RM A9101 /.2 i 3 / /.0 “n n L1 u 04 .6 I/ A 1 I / H C12 /- / m /1 00 .4 H V-4 .2 0 -.2 I I ., Y 1/ -f.o -8 –4 ‘4 0 Angle of attock, u, tfeg 8 /2 [f) M, 0.90. Figure 1o.-- Continued. for ~ =4” , Figure /0.- Continued. . 45 NACA RMA9101 . *. b 04 t--l .4 u .2 A ‘Z v -4 D -6 H I I — ..- v -/0 , I I ‘.8~ -4 ‘8 Angle 0 of dtuck,”a, I I 4 1 L_ 8 deg (h) M, 0.95. Figure 10.– Concluded. for 4=4” 46 NACA RM A9101 . . 10 .8 .6 .4 & ~’ .2 .$ 80 g 4 -.2 -.4 . — -.6 -.8 0 .04 .08 .12 .16 Dog .20 .24 coefficient, .28 .32 for ~ ‘ 4° Co (U)M, O.20. Figure //.– The drug charucterisilcs of the airplane model mounted in the extended whg-chord plane. R, 2, QUO,000. c -——. with the horizontal toil NACARM A9101 47 ‘ . . /.,0 . 8 6 ,4 ,2 ‘o ~ 4 -Vttttli” -.2 -.4 -.6 -8.- 0 .04 Drag .08 ./2 coeffkien~ C’ ./6 .20 . [b) M, 0.50. Figure Il.— Continued. .24 .28 for 4=4 4a NACA RM A9101 . /.0 .8 .6 .4 6’ S* .2 .? f. $ 4 “.2 -.4 -.6 .— -.8 0 .04 .08 ./2 Drug coefficient, ./6 .20 .24 .28 .32 for ~= 4° 6“ (C)M, O.70. Figure Il.– Gontinued. . 49 NACA RM A9101 /.2 10 .8 .6. .4 e ~c .: .2 .% k Q Q ‘o ~ 4 -.2 -.4 -.6 -.8 .04 0 .08 Drag ./2 .20 .16 coefficient GD (@M, Figure . //.– Continued . . .24 0.80. .28 for lj “ 4° 50 NACA BM A9101 —---- , . AZ /.0 .8 .6 .4 ‘~- .2 .% fo ● Q v .$ -,2 4 -.4 -.6 -.80 .04 .08 ./2 ./6 .20 .24 (e) M, 0.85. Figure //.— Coty’inueo! _ ........ . . . . . .28 for I)= 4° 51 NACA RM A9101 /.2 /.0 .8 .6 .4 ! I Ua V4 .- v \ .1 \ * \ \ 0 .04 .08 Dreg coefficient .20 Go figure II.– Continued, for 7 II I)*4” ... , (f) M, . . J6 ./2 ,“ 0.90. NACARM A9101 ● . a “ Figure //.– Continued. 53 NACA RM A9101 . . * . . .04 o .08 Dug //.– Goncluded. ./6 coefficient, (h)M, . Figure ./2 CD 0.95. for ifs4” .6 .5 .4 .3 .2 ./ O :/ (WM, 12.- The pitching- horizonkd toil moment mounted in the Wended . :5 :6 ;7 Cm ii! of wing-chord . -.4 0.20. characteristics , :3 metficieti, FVching+noment Figure :2 the airpkme pkme. model with the R, 2,00~000. . . ✌ ✎ ● ☛ /0 .8 , I .6 f’1 A J/ r . ~,+’ . H++ .4 H-ii—t—H—l .2 i ,, ! I i .1 .6 .5 .4 .3 2 ./ Pitching-moment (b)M, Figure 12.- I I I I I I -.8 I I Continued. o ./ -2 c~ftkien~ :3 4 ?5 I I I I -6 I i ?7 Cm a 50. U U ul m m .8 .6 @ .4 g \ .2 1 0 Ill 5 -.2 -.4 -.6 -.8 .6 .5 .4 .3 .2 J O Rkhing+noment :1 coeftM@ ?2 73 -.4 75 -6 ~ (c) Aft 0.70. figure /2.– Continued. 1 . 1 . , # a /.2 g o P ,8 .6 d -.2 -.4 d II I I r 1/ r I -.6 w Fiiiiiiiiiil 1 1 / I I 6 Wtrwt 1 I 1 I 1 1 I I I I I 1 1 1 1 I I I I I A 1 1 tum UJ r 1 -.8’ 7/ .54.32]0 fYfdi~+xwhm t :2 cdfuk?n~ (d) Ai, 0.80. figure 12Y Continaf. -3 G -# -5 76 :7 . .- i.2 /.0 .8 .6 e .4 + $ .5 Q g -2 o 4 -.2 -4 -.6 -B .63” .4.3.2./0 :1 IWcbihg-monkmf {e) IV, figure . /2.- co&ihi”m~ 72 -3 -.4 :5 -.6 5! G 0.85. Continued. 1 ✌ ● ‘ [2 /,0 .8 .6 .4 t+ ,2 . $ “$ 0 % $ -.2 -.4 -,6 -.8 -lo ...6 ,5 ,4 .3 .2 J O + c06mim~ R7dlihg-mment (f) i7gure i2,– Continued. M, 0.9Q 72 ~ .3 + .5 d 33 OH I , IQ I I 1 f I II I I I A I t , “tb!tkki 76 / /? v / / x v k r P f / / 3 4 $ 1 Jlllliil / -J3- .6 .5 4 .3 .2 ,/ RwVhg—mamn (9) ~, Rgufl /2.– I I I 1 I I L/ Continued. . o t -] cv#%bkw~ 0.92. :2 G -3 4 -5 * . figore 12.- Concluded. 62 NACA RM A9101 . .— . . .- .8 .6 .4 w .2i R w . o .s -.2 / / / -.4 =4s=’ c‘ ‘.6 -8 -4 0 48 /2 Angle d (a)R’, Figure 13– The lift characteristics mounted in the extended dn, 3(7; ~f ,50° ; M, 0.20. uttuck, /6 20 for ~= 4* a, deg 2,000,000. of the airplane model wing-chord plane and with the horizom’ul with the flaps deflected. tall ..”” “., .:., NAC!ARM A9101 Angle of attack, a,a’t?g (’lot?, q ● figure . . 13? Continued. 000,000. 63 64 NACA RM A9101 . II H u /.0 . %- 04 ❑ 12 00 II I .6U Q -8 I I _,/’f p J-( ~/ 4%4 .4 . .2 . 0 . ‘.2 -.4 ‘L ‘~L 1/ 1/ y, I@ f T‘ / c, ? <( 1 -.6 -8 -4 0 4 8 Angle of ottoc~ a,deg (C)R, Figure 13.– 1 I /2 I I I 16 10,000,000. Concluded. @ya?%wEIi? -.,.-. . .... I I for ~=4” I I I I I NACA RM A9101 65 . ● L6 L4 12 /.0 .8 . G’ ~- .6 .= Q s~ .4 Q s .2 0 -.2 -.4 -.6 0 .04 .08 ./2 .20 ./6 Drag coefficien~ .24 Go .28 .32 for 1;= 4* (0)t?, 2,000,000. Figure 14.– The drag characteristics . mounted 8f, . in the extended 50°; M, 0.20. of the airplane wing-chord plane model with the horizom’al with the flaps deflected. & 30°; tail 66 HACA RM A9101 1.(2 .8 . . 0 .04 .08 ./2 ./6 .20 .24 .28 Drag coeffickm~ Go (b) R, 6,000,000. Figure 14- Continued. .32 for t) = 4“ NACARM 67 A9101 a . . 0 .04 .08 .12 .~6 20 Drug coefflcien$ .24 .28 Co (c) R, 10,000,000. Figure . . /4. - Concluded. for~=4° 68 NACA RM A9101 -4 , . . -i ? . .6 .5 .4 .3 .2’ ./ Pi%%ing-moment O + coefficiii~ :2 -3 -.4 -5 ~ (0) R, 2,000,000 Figure /5. - 7he pitching-moment chffrocteristics of the oirplune model with the horizontal tai/ mounted in the Wended wing-chord plane and with the f/aps deflected ~, 30~ +, 50~ M, 0.20. -.— .......--.. W“d 69 NACA RM A9101 . .6 .5 4 .3 ?2 .2 J O J Pitching-moment co~cb~ (& (b) R, ~000,000. Figure 15F Gontinued. ?3 -.4 ?5 ;6 NACA RM A9101 70 . .- .- --. .Q* . b o 4 “S ❑ 2 o“ -2 ~ %. 4 0 .4 A -6 4 -8 v -/0 ! D .2 ,n . , /’%for.zontul r }1’1-l 0 .— -4 v El b . toil off -.2 -.61 ‘ ‘“ .6 Figure .5 /5.– I I I I .4 Concluded. I .3 I 1 I , I 1 1 ! 1 1 72 + .2 .1 0 Waling-moment coatficien~ C% (c) R, /O,OO(?,000. 1 .. I :3 -.4 75 71 “ NACA RM A9101 . /.0 M .8 — n o o,5~ 0.70 A 0.80 D 0.90 F — a 0.92 v 0.95 .4 6 -~ c .-a! u .2 c g . . . o g 4 . . T /‘ -.2 > , L. / / 77 14 ) V / ‘.6 TP ‘< ‘F . -B ‘8 -4 0 4 Angle Figure for ‘ 16.— ~he lift choracferisiics moum’lng the horizontal fail /2 8 of oftuc~ a, of the ubove the wing /6 20 for IW= 0.20 o’eg and fuselage. fuselage with R, 4000,000. ?%e bracket NACA R’MA9101 * . . . .“ . b Figure the .04 .08 ,/2 ./6 ,20 ,24 Bag coeflicien~ Co /7. - The drffg L2$ff rueteristics bracket R, 2,000,000. for mounting ,28 for M= 0.20 of the wing ond fuseloge the horizontal with toil ubove the fuseluge. . — “.. NACARM 73 A9101 /. o [ Illllllld I I I I I I m I I I Islll m .8 r / \ -. I&l .w .6 < IY I I\ 1 rJ I .4 \ \ I /1 141 d /1 w II D / I IY-HHI(I I \ I \ I w s I II 121 A i I Ihf 00.20 I •10.50 \ 00.70 A 0.80 \ V 0.85 ~ \I ., I I . . -.2 ‘. 6 ‘.8 -./ o J for Pitching-moment coefflclent, Figure /8.-The pitching-moment fuselage . . wlfh fh8 the fuselage, bracket #?t 2,000,000. M=O.20 CM characteristics of the wing for mounting the horlzontol uno’ fuil above .8 .6 -.4 -.6 0 .04 .08 Dug .E .16 CO#w@ .~ CD .24 .28 -8 ..3? -4 .36 0 Am of . .40 4 a~ck, 8 /2 a, dq E 20 ./ o -.1 Fh’chlng —nwmew Figure j9.mounting . . Zhe Iift,drag, the horizontal and pitching toil above ‘momwt charoc~e~~tics the fuselage, . flaps of deffedeo! the m“W O@ + , 30~ {, fUSehI@ 5fl wifh the M, 0.20. . . ?2 co&%4m~ bracket (& for E! g t-i o 1- NACA RM A9101 r . /.0 .8 deg 04 ❑ .6 ‘t 2 00 A ‘2 V4 & .4 D H -6 . .i ‘8 0 -4 Angle of otfoc~ 4 GZ,deg (’U)M, Figure 20.fail ~he lift mounted characteristics above the fuselage. . 8 8 /6 /2 for tj*4” 0.20. of the ui~lune R, 2,000,000. modei with the horizontal . . 76 a!MaEEe) .. ..-.. . NACA RM A910L . . . (b) M, 0.50, Figure 20.– Gontinued . . NACARM A9101 77 /.0 . ‘t deg .8 — 04 00 .6 — A (j ‘z v-4 .4 — P ‘6 / (J // / 4 / t .2 )$ /!/ * o -.2 z. \- 1 $- F~ ● ( J( ~ <> -.4 c) r L ) ~ , c <( -.6 L> Kr { D‘ ~– -.8 + -4 0 4 Angle /2 /6 8 of otfuc~ U, deg (c) M, Figure 20.— 0.70. Continued. 9 for it=4* 78 NACARM A9101 . . /.0 -, ~, ,8 — deg U2 .6 “ 00 A ‘2 .4 — G ~.? u s w ~ V4 .2 0 $ 4 -.2 . . . ‘.8 ‘8 + o 4 /2 8 /6 Angle of ottac~ u, deg (d)M, Figure 20.— Gontinued. 0.80. for #j =4° 79 NACA RM A9101 ‘ . /.0 . ‘f , ? deg 8— 04 ❑ .6 — z 3 00 / A ‘2 -, c.5—- “TtttT7m -.8 FF# -/.0 ~ ‘8 I –4 I I I I o 4 Angle of uttac~ 8 /2 a, deg (e] M, 0.85. Figu re . . 20.— Gontinued /6 for ~=4* 80 NACARM A9101 . /.0 . ‘t deg .8 — 04 R2 00 .6— ‘A ‘2 V4 .4— D -6 ..— ‘8 0 -4 Angle of attack, 4 a, 8 deg {f) M, Figure 20.– Continued. 0.90. /2 for 4=4” . NACARM 81 A9101 ., 0 -4 ‘8 Angle of ufftick, 4 a, &g (g) M, 0.92. , Figure -. 20.— Continued. 8 for it=4” &? ...’-...- - NACA RM A9101 . — . . * Figure 20. – Concluded. . NACA RM A9101 83 . . Drug coefficien~ ~ (a)M, 0.20. Figure 2L— The drag of the uirpkme model wii’h the R, 2,000,000. chorucieristics mounted ubove the fuseluge. horizontal fail NACARMA9101 84 . . , . -.UO .04 .08 ./2 .16 N.11”1111 .20 Dreg coefficient, (b) M, 0.50. Figure 21.- Continued. ~i .24 Go .28 A for & = 4° I NACA RM A9101 (c)M, Figure 2/.– Continued. 0.70. HACA FM A9101 86 /.0 I .8 -. .6 .4 ~ -1 -.2 -.4 I\. -.6 .-. 0 .04 .08 ./2 ./6 I .20 I I .24 Drag coefficien~ (d)Af, 0.80. Figure 2/.– I .28 I I fw&4° Go , Continued. . . NACA RM A9101 , , 1.0 .8 / / .6 j / deg / / Y .4 / , ? # / k .2 ● . $ .9 “2 $ Q ; 4 4 r f 4 J‘ 4 ❑ 2 00 A ‘2 / w “o f v -4 J o -.2 \ . -.4 — Y — -.6 ‘.8 -/.0 o .04 .08 ./2 ./6 Drag coefficien~ .20 Go (c)M, Figure 2L- . Gontinued. 0.85. .24 .28 for @4° 88 NACARM A9101 . # . .“ -. — . . NACA , . . . . . Dreg coefficient, CD (9) M, 0.92. Figure 21.- Continued. ● .. .. . NACA RM A9101 9 . . P / I A () \ \ I I \ . -.4 . o .04 Bag .08 coefficient, ./2 G~ W figure 2/.— Cone/uded. ./6 M, 0.95. for ~S4e . . Ill Pitching-moment Figure the 22.— The fuseloge, pitching-roomed R, 2,000,000. characteristics of coefficient {0) M, the air.okw Gm 0.20. modei with the imrizonfui taii mounted . — 1 1 + .8 04 H 1 1 I I I I , I I I t I , , I , 1 , deg H .6 H 1 I ❑ I I z 00 A ‘2 ,4 ,, I 1/ I ) ~i 1 I 1 Ciii :. i ! ,, !. ,. ; L I -.4 -.6 , –.0 .6 .5 # .3 .2 1 t7tii@—momen 0 22? :2 -3 + s s -5 t cQIS%ficim~ l% (b) M, 0.50. Figure 7/ ‘!2 &H“ 0 I-J Continued. * ., , . , ,-* !$! UY u U3 Lo .F .8 .6 .4 e ~s .2 ~ * lo g 4 -.2 -.4 -.6 -.8 .5.4 .3.P 71 ’JO Pitcting-nmment coefflcjen~ (d)hf, figure . . 22.– %? :3 -.4 + ~ 0.80. Continued , ,,, . .1 Ill!!l ,7 .6 .5 .4 .3 .2 J FrMiw?wmw fe)M, Figun3 22.— Continued. t 7/ O meftkkq 0.85. 72 G) -3 -.4 75 76 II!l”l!!l!l!bll? j .8 H-i I I o’eg m o 4 ❑ 2 I I I I I l#T I <1 I !/ 8 r II/ ‘< .4 -.6 -.8 -10 .7 .6 .5 .4 .3 .2 J Pltchihg-nmment (f) Figure ., 22,- Conflnued. M, 0.90 O -./ coefficient -.2 Cm -3 -.4 -.5 .8 .6 .4 -.6 -.8 -10 .7 .6 ,5 # .3 .2 .1 Rhhing+mment (g)/U, Figure 22T Continued. 0 cafXWn~ 0.92. ‘~ ‘. & ?3 -.4 “5 . ,71 :,, ,. ., S6 /1 J 1/1 v l.?! 1 ‘“2 I&. .C 1, . I Pitching-moment coefficient, I I I Cm !2! (w IW,O.95. Figure I 22.-Concluded. 1 99 NACA RM A9101 . . M O ❑ .8 0 A 0.20, 0.20, 0.92, Originul Afodifie’d Original fuselage fuselage fuselage 0.93, Modified fuselage .6 f .4 (3’ ~w ~ .% Q $ m Q . Q ~ 4 > .2 4 o < }$ -.2 ( / . P -.4 / { ~$ J $ { / v / / ‘.6 +-+V- 4Q ‘.8 .3 o ./ .2 Pifchingwnoment 4 .3 -f :2 coefficien~ Cm .2 ./ :3 o Pitching-moment Figure 23.- The effect on the pitching-moment -1 ‘ coefficient characteristics fuselage of the airplane model with the horizontal /?,2,000, 000; $,00. -.4 72 :3 Gm of modifying the toil obove the fuselage. 100 XAC!ARM A9101 . 7,4 4 I 1 1 1 4 1 /.2 Q- “- rr deg o d{ 0 h 1 i I? 1q -.2 * / f +\” n 2,000,000 4 D 6,000,000 4 0 /0,000,000 4 A 2,000,000 0 v 6,000,000 0 @ /0,000,000 o 4 2,000,000 -8 v 6,000,000 -8 v 10,000,000 -8 1 * . . I -.4 t-H t 71TTtt-1 L ‘.6 -8 0 -4 Angle Figure 24- mounted M, 0.20. The lift ab we 4 8 of attack, characteristics the fuselage /2 16 20 for ~ = 4“ a, &g of i’he airplane model and with the flaps with the horizontal o’eflecfeo! an, 30°; ~f 50°; Iaij 1 . 101 NACA RM A9101 0 -.2 -.4 -. 6 0 .04 .08 ./2 ./6 ,20 Drug Figure 25.- 8f, 50: obove Iul 0.20. .28 .32 .36 for~=4° coeffickm~ CD The drag churacferistics tail mounted .24 of the uirplone model with the horizontal the fuse loge and with the flops deflecfed #n,30°: - 102 NACA FM A9101 . . — ,..& -- ‘... .* .* ,.!. . .. :-. — .7 .6 .5 4 .3 .2 Pitching+wnmot J O coefickw~ -J ?2 -3 -.4 ~ of the oirplone model with the 26.- The pitching -mome@ characteristics horizontal tail niounted ubove the fuseluge and with the flops deflected, Figure 4,30; 4,50°; M, 0.20. -5 , 1 H 1 1 I ‘Q . .~~ -0-- ~oxjmum ~s~ -!3-- Wake boundaries ---— Horizontof -80 tail ‘. \ —Horizontal ‘j in the e.sh?ndad wing-chwd — \ [email protected] tad abova ##one the ‘x, •. \ ‘\ 1 \ I fusebge +(x)’ 16-8 -8404812 Angle of dock, a, 4 O@ 0 Angie 2Z - The variation horizontal-tail positkms with Ongie of attack with reference 4 c# dto~ I I 8 L? 16 a, deg (b) M, 0.50. (a) M. 0.20. Figure I of the wake-boundary, to the wing-chord pfane maxiwm-pressure-fess d a =OO. Wing and locations fuselage. ond R, 2,000,000. -. ?-4 8 -8-404 Angle of otfad, 12 B 0 /@71e of aftock, 8 u, /2 deg U, d~ 16 -84 Angle (~M, (’C)IW, 070. Figure 4 &280. . 0 4 8f216~B ofottock, u“, deg (e)tW, 0.85 v 27.-i2xv’inued. . . ,1’!’! 1’ . , E! I\, , fuselage \\ \\ \ \ \ 1 \ A \ E -7 \ + 1’ ‘ \ \ \lz \ “ I ~\ \ \ \ Y \“ \ / 1 .1 \ \ \ r \ H--H — ‘8 An@ Figure27 .– of aftoc~ (f) .44,0.90. Conclu&d. a, d?p 0 -4 Angh -=z?@-2 d (d 4 ottook, u, &g M ,0.95. 8 106 NACARM A9101 . 80 60 Eli \ ~ R 2,000,000 #axijnum pressure /0ss ~ 6,000,000 Maximum pressure loss *108000,000 ~ 28000,000 Mt3ximumpressure /0ss Wake boundaries ~ 6,000,000 ~/0,000, 000 Wake boundaries Woke boundaries . . . ‘8 -4 0 4 8 /2 16 20 Angle of ottack, U, o’eg Figure 28. - The variation with ongle of uttock of the woke-boundary, maximum-pressure-loss locations ond horizQnfal- tail positions with reference to the wing-chord plane at u” O‘. Wing ond fuselage with the flaps deflected %n,30~ %t,50~ M, 0.20. NACARM 107 A9101 .241 0 Wing ulone 1 .20 ./6 ./2 “ . . .08 .04 0 , /0 Disfonce a..ove wing -chord planeat Flgufe 29. - Wing woke semitdl areu. ~= O”, percenf ct patterns at the cenfroht Q, 61 M, 0.85; R, 2,000,000. . . of 108 NACA RM A9101 . /,00 .98 .96 .94 .92 .90 .88 .86 .84 o .3 * FFFF -6 ‘4 -2 (o) Horizonfu/ 0 P 4 6 8 10 toil mounted in the extended wing-chord plone. /2 , Figure 30.- The variation with ong/a of ottuok of the rotio of the dynumic pressure at the Iuil to free -stream dynamic pressure. R, 2,000,000. NACA RM A9101 log a . . /.00 / .98 \ l\\ \ .96 \ .94 \\\ , 1 .92 . .90 . .88 .86 ‘6 -4 -2 Angle 0 of 2 6 4 affoc~a,okg (b} Horizon?oi foil mounted obove the fuseluge. Figure 30.- Concluded. . 8 10 /2 c 110 NACA RM A9101 . . . . Figure 3/ .- The variution pressure & 300;+, ut the 50: fail with ungle of attack to free -streum of the rutio dynumic M, 0.20. -QzEis2$ pressure. of the dynamic Flups deflected. NACARM 111 A9101 ● .90 ● .86 .82 . .78 1 .74 .70 .66 .62 . .58 .54 ti .50 I . I I 1 1 1 I1 1 1 1 1 m llb&Ll_lll I I I 1 I I I 1 I I I I I I I Iu=o.50 / - .46 . .42 .38 . . . .34 .30 .26 .22 H+H+H+t-+t-tti I -- A4=0.20 .18 -8-6-4-20246 8 /0 Angle of ai’fuck, a, deg (0) Horizontal Figure 32.- d attack. tail mounted in the extended wing-chord phne. The variation of Mach number at the toil with angle 112 NACA NM A9101 . .98 A4=o.95 .94 . \ M=O.92 ~.o.go .90 — “ MnO.85 .86 .82 M=O.80 k .78 .74 -. -. .70 ~.o. TT-Kr” 70 I 1 .66 ● .62 ✎✌ .52 M=O.50 .48 \ .44 .24 .20 ‘ \ M=o.zo v — ./6 -6 -4 -2 Angle (b) Figure 32.- Horizonful Concluded. 0 of attock, 2 a, toil mounted 4 6 8 o%g above the fuseloge. 10 113 NAC!ARM A9101 / / . I L4- I -41 ‘8 (0) Hori20nt01 Figure 33. - angle centroid of 1.4’ I -4 Angle toil I of uffffc~ mounted down wush at semitoil i I 0 The voriation of the I in’ the 80 *0. ‘ ml I I I 8 4 I /2 a, o’eg extendeti with angle o position wing-chord of attack of the corresponding oreu. R, 2,000,000. plone. effective to the 114 NACA RM A9101 . “ ‘ ‘ ~ / ~---- ---- .#-- ----- I -;=O.sO - / /‘ . / M“O.20 #---.///- --I ~“ 0.70 I ‘ I M“O.80 , ““’ #“ g . . . ‘8 -4 0 Angle figure of attack, toil mounted (6) Hort’zonful 33. -Concluded. 4 a, 8 /2 f~e fuseloge, deg ti&oVe . . .. \.._.., ~z!E!3b . 115 liACARM A9101 ‘2 [ R -2,000,000 ----6,000,000 —+0,000,000 . 8 .—. ---A 4 0 4 (a) ‘8 -4 Horizontal Angle of attack, e,deg toil mounted In the extended 8 /2 wing-chord plane. 4 0 -4 -8 4 0 8 /2 Angle of attack, a, deg (b) Horizontal Figure 34?tfownwash area. Flops The variation of oposition de ffected. tail mounted above the fuselage. with angle of utfuck corresponding &,30e; c?f,50”; to the of the effective cenfroldof M90.Z0. the ongle of semifuil .. .. . ,, (o) (2 ./ Horkontol fail mounted .3 .2 in the .4 .5 number, Moth (b) Figure 35.- The variation Wth Horikon?oi Mach number tail of mounted the . extendd .6 efficiency . @ne. .7 .8 .9 Lo M obove tail wing-chord fuselage. factor. * NACA RM A9101 . . /.0 ‘ . R — 2,000,000 -----6,000,000 — —/0,000,000 .9 ‘ v qyq .8 .7 ~a) Horizontal fail mounfed in the exfended wing-chord plone. .9 - / - -. ----- --- \ t- *.M “ .8 v ‘Yq ‘ .e@H —--. - : /~ -– .7 ‘8 -4 .4 0 Angle of offack, a, deg 8 /2 lb) Horizontal fail mounfed above the fuselage. Figure 36.- The variation Flaps de flecfed, {,30~+, of fhe fail efficiency 50~M,0.20. factor with angle ofatfock. -. .- 8 ~ s0.6 . -. ‘ —. . - --- — . -.. ---- _ - — . i o,~ - 0.3 . —_ -. . -. --. _ . . _ _ : -. -- = a . — 0.2 . __ of ---– -:-.–--—— -----.-—-- ----------------....----- —----..---_ _... _._. ___ . ~.- 0 8 .- s CL 0“6 --- .- -. . ;;4 -- -- -- - -- ‘- - -- -- 4 I--H-’”zr 1 1 of J “o ---- ------ .4? ---- ------ Mach Figure 37.— coefflclenfs. The varivtion R, 2,00i9,00~ with tail Mach it, 0? --- .4 ..3 [b) Horizontal --- ‘-*’- WWUIJ I uu ’-J number ---- -_ uuuw ‘A-”’- of the ----- --- \ . --- / -- -. — -, ---- ---- —--— ------ .7 .6 .5 number, II --- ---- .9 .8 i.o M trtw ‘k- angle usaruyu. ‘--I--- = J of attack of the airplane model for various Iif t E El & H o P E! . Mach (b) Figure 38.- ?%e variation coefficients. Horizontal tail with Mach R, 2, 000,000; mounted n:mber ~, O. of number, above the drag M the fuselage, coefficient of the airplane model fir various lift -- kj ii 8 Wlforizontol tail mounted in the extended wing-chord plane. J ----- y G“ “ ;:? .0 . -------” -----” ----” ‘---- “--=-- “---- ----- ‘“--- “-- ~ ~o ___ r 0.4 ‘- - ,---- -- //. -- 0.3 Q u ----- - “— ‘-- — ~ ~ --- / ‘- . .. . . Y “ Y, . ‘.. , \ 3 $-.1 _ -. _- 0.6 ---- 1 ~ ~ .. K =&=– -.2 o .2 ./ .3 Mach (b) Figure 39: airpfune t., .4 .3 The variation modet for Horizontal tail with number varibus Moth lift mounted coefficients. r .6 number, , of the .7 ,8 .9 Lo M above the fiselage. pitching-moment R, 2,000,00~ ~, O: coefficient of jhe . . $?4 II - b’ 4s’ 9 I .2 ---- ----- (a) Horizontal Illil _ - - Q,6 I ----- ----- toil mounted ----- ---- ---- in the ---- I I ------------ extended I I --- wing-chord I —-/---- I “- 1 1 plane. . / CL*.6 .4 .2 ---- 0, --- --— ---- ---- ---- ____ ---- ---- ---- -- - ----- ---- ~ ---- ----- J \ o o .2 J .3 .4 .5 .6 Mach (b) Horizontal Figure 40.- The mriotion corresponding to the with Mach centroid number of the semitoil number, tail mounted of the ores .7 .9 lo M above effective for .8 various the fuselage, angle lift of downwosh coefficients, at a position R, 2,000,000. ~ i’- Iv w 1.0 .8 !!. . HorizontoJ toil mounted in the extended ~ ~ ~ - Horiz ----- ontal ----- toil moun ted above .- —- ----------- I I I - ----- 1 plane ~ ~ ~ wfng-chord ~ ~ ----the fusela ----- go ---- ---- .-:-- -a-- I I [ I 0 0 I 1 I .3 .t? J I I I .4 .5 Mach Figure 4/ .– ifownwosh The with variation angle with of Mach attuck. number I of I .6 number, /he I rate I .7 .8 ‘-- I 1 I I I .9 1 I 10 M of change of the effective angle ~ of E$! g 1+ R, 2,000,000 ~ ,. , , NACARM 123 A9101 .8 ,. Qti \ /“ .6 4=-4; __ ----------- --—---------------3 ‘-- ---/ / .4 ,/ -2° s .; .2 Q “w k :0 Q ~ ~ -.2 -/ 0 - 0° ‘- - \ / ‘ . \ / 4 / u 6 / ~ ‘- K ‘-- ‘- - \ 1,-) -.4 -.6 ~a) Horlzonfal tailmounted . intheexfended wing-chord p/a;e. .8 . iL-40 t --3& -2° .6 ● (y .4 * ~. -= q ~----___---- --------- ---- ----- --- _&, I ‘k . ‘ -I I I 0° 1 1 - $ ‘l-t-t-t’: 0 I ~ .2 I 1 1 1 1 - 1 I , — / , “ 1 b-l —- l— ~. 4°~ I ‘--‘-- ‘- ‘-“ ‘-- 2L , r \ -- — - --. I I . Y % / \ \ \ ,I .4 I I \ / I =?3= .6 0 J .2 .3 .4 .5 Much number, @} Horizontoltoil mounfed .6 .7 .8 .9 lo Af a&ovefhe fuseioge. Figure 42.- The vorjation of fhe /jft Coefficient for Cm=O with Mach number. R, gooqooo. .- -. .8 \ \ \ \ \ \ rl=4 ~,. b, .6 . 4 II .2 0 .5 --H--H-W ,6 .7 .8 Mwh Figure wing 43. - The lift loading an altitude ,. of of coefficient 100 pounds 10,000 number requt’red per .9 10 ~ M for square an airplane foot in fli&f with a at feet. . . . , o z! b % + -2 * $ ~ -4 (o) Horizonto/ toil mounted in the ~Ambd wi~-chord plane. Z 4 k % * a 32 \ o * *O v -2 -4 0 .1 .2 .3 ,5 .4 Abchnmhw, (b) ~igtue 44. - Variation h flight Ot laOOO with Hor’izontol Mach feet. number Wing 100ding the 100 .8 .7 10 .9 M toll mounted of .6 angle rends above the fuseioge. of stobillzer Per Wu~e seftlng f~. to balance Genter of 9fovi!Y an olrpione ot O“25c’= .- .8 c? .4 -.4 th)M, (NM, a2t2 {i#M. 0.50. 0.80. .8 .6 -.4- 8 -4 0 4 -8-404 -8-404 Elevator fb)hf, 0.85. Figure 45. - 7i?e lift characteristics elemnbr. . . (f)hf, deflecfion,~, 0.90. of the imlotd R, ~O@700. ., -8404 deg @M, 0.92. toil with a 20-percent-area 6)M,0.94. cmsfant-chcwd ,. . ‘. . , --- f)=o .. ~ L ~- . /“ # /“ / . . - / / / G-z .6; c=== ,~ , — -. /7.84 ( / I ( n=3 {o) Horizontal mounted in the extended /). I ~ ?! w toil -2 wing-c40rd I 1 1 I I I ..-++- plane. 1 I *- {, O: 1 Y kt=l I I -4 -6 0 .1 .2 ,6 .3 Ah (b) Horizontal Figure 46.- at 10,000 Voriotion tith Moth feet. Wing ioading number ms&r, toil mounted of JOO pounds the e/evator per squore .7 ,8 1,0 .9 M ofwve the fuselage. deflection foot. Center to bolance of gravity +,2: an airplane at 0.25 c! in f/ight —. I u “ill 03 .L 4 .4ngle of stabilizer Figure 47.- The angle the horizontal ofstobifizer tail mounted 0 -4 0 sefting,~, setting in tie .04 deg for extended .08 . GO and wi~-chord iWg Me .12 coefficient, corresponding plane .16 and the .24 .20 CD drag flaps coefficient deflected. for ~, 30°; modef the ~, 50°; witi &l-l M, O.20. 0 1- .. . + . , v , . II -2 0 Lift Figure in I 48. the – The extended lift-drug wing-chord rqtio .6 .4 .8? coefficient, of the, pfune. .8 /.4 CL airplane 8n, 30°; model with the horizontal ~, 50?j (%, O ; M, 0.20. tail mounted 130 l?ACARM A910~ , ? -+ 160 -1-L2rL 200 240 Gliding speed, mph . 280 Figure 49.- Variation with gliding speed of the sinking qveed of on airphne in power-off flight at sea level. Wing loading 100pounds per squure foot. . Center of gravity at O.~Sc! ~ 1 b “, 9 ._ . 4 NACA-Langley -12-8-49- S’75