E I RESEARCH MEMORANDUM AERODYNAMIC CBAFLACTERISTICS OF A WING WITH QUARTER-CHOFCO I LINF, SWEPT BACK No,ASPECT RATIO 4, TAPER i FLAT10 0.6,AND NACA 65A006 AIRFOIL SECTION TRANSONIC-BUMP METHOD By Thomas f. .Erg;Jr., and Boyd C. Myers, II Langley Aeronautical Laboratory Langley Air Force Base, Va. """- .NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON September 6, 1949 UNCUSSiFfg-? P a RATIO 0.6,. e NAGA 6s006 -OIL SF,CTkON t . AB part of a transonic resemch.prog??am, a series of win@Oay combinations qe being investigated in the Langley high-peed 7- by 1o"foot' tunnel over a Mach nmiber range from about 0.60 to 1.18, by use of the tranaonic4Ump test technique. i This paper presents the results-of the investigation of a win-ane and a wing-frmelage configuration amploykg a wing with quartemhard line swept back 600, aspect ratio 4, taper ratio 0.6, and an WA65~006 airfpii section. The results are presentedas lift, drag, pitchine moment, and bendineanent coefficients for'both confipatim. The effect of a wing fence on the wing-fuselage charsbteristics was also investigated. In addition,e.ffective downwaah angles and point dynamic 'pressures for a range of tail heights at a probable tail .length &e presented. for the two configuratiomFnvestlgated. Only a brief ' Wsie is given in order to facilitate the publiehhg of the data. IIJTRODUCTION A series of wiwfuselage canbinations ISbeing inveetigated in the Langley high-peed 7- by 10-foot tunnel to study the effects of,- gemtry on longitudinal stability characteristic8 at transonic speeda. In the trm.8dc-bmq technique used, a Mach nmber range of 0.60 to 1.18 is obtained.Previous data published in this series are presented in references 1 to 3. 2 - XACA RM ~9~27 This paper yresents the resulk8 of the investigation of the whg- .. alone and WTng-fuselage configurations .employing a wing with the qumter-chord line sweptback 60°, aspect ratio 4, taper ratio 0.6, arid t an NACA 65~006airfoil section parallel to the free stream. The wing of the semiepan model had 60° of sweepback referred to the quarter-chord line, aspect ratio 4, taper ratio 0.6, asd an MACA 65~006 airfoil section (reference 4) parallel to the free stream. The wing was made of steel and the fuselage of brms. A hrwview drawing of the model is presented in figure I and ordhates of the fwelage of fineness ratlo 10 are given in table I. Details of a wing-fence arrangement tested are shown in figure 2. The model was mounted On an electrical'straqn-gage balance enclosed in the bump, asd the lift, drag, pitching moment, and bending moment about the model plane of symmetry were mea8ured with calibrated potentiometers. Effective dowriwmh angles were detemed far a range of tail heights by measuring the floatfng angles opthe tails at five different positions with calibrated slide"wire potentfameters.Details of the floating tails are given in figures 2 and 3, while a view of the model mounted on the bmp, showing three of the floating tails, fa given in figure 4. The tails used in %hi8 investigation me the 8- a,a those used in references 1 to 3. A total-pressure rake W&B used to determine the dynamic+resame ratios for a range of tail heights in a plane whlch contained the 2-ercentmean+rodynamic"chord point of-the free-floating tala. The total-pessure tubes *e spaced 1/8 inch apart near the chord line exte-ded and 1/4 inch ap& asewhere. I 3 . c, I bendlnmnt coefficient about root chord Use (at plane I 9 effective dsnamfc .pressure over span of mdel, po&s per square foot (p~2/2) S twice wing mea of semiapan model, 0.125 squaze foot - C mean aerodgnamic chord of wing, 0.181 foot; based on relationship c%y. (us% the theoreticaltip) S .. C local wing chord - .I Ct mean aerodynamic chord of tail b twice span of semispan del spanwise distance fram wing root -. P air density, slugs per cubic foot v free-stream velocity, feet per second M effective Mach number over span of model . Mach M2 local number I Ma average chordwise local Mach nurmber - R Reynolds number of wing based an c . a angle .of attack, degrees 1 - .E effective downwash angle, de@;rees ratio of point dynamic pressure, along a line c.mta5ning the ! Q qumter-chord points of the mean aerodynamic chords of the free-floating tails, to the local free+tream.dynamic preseure " 4 - NACA RM LgG27 ht tail heightrelative'to wing chord plane extended, perced semispan; positive for tail positions above chord plane extended I a.c. aerodynamic center Subscripts: M at con8th-t Mach nmber CL = 0 at zero lift The tests were conducted in the Langley high-peed 7- by IO-foot tmel by use of an adaptation of the XWCA uin@low technique far obtaining transonic speeds. Themethod used involve6 the mounting of a model in the high-velocity flow field generated over the curved surface of a. bmrg locatea on the tunnel floor. (see reference 5.) Typical contours of 1ocal"ach nunibere in the region of the model location orl the bump, obtained from surveys with no model in posttion, are shown in figure 5. There ia a Mach number gradient whioh results in a diffQrence af about 0.04 over the'span of the model at the lowest aid higheet .Mach numbers with a maximum difference of about 0.07 present at a Machnumber of about 1.0. The chordwise Machnumber variation is generally lepe than 0.01. Ro.attampt has been made to evaluate the effects of these spanwise and chordwise variations in Mach number.The long-daafied lines shown neax the wlng.root repreeent a locd Mach number 5 percent below the rmxlmm value and indicate the extent of the bmrg boundary layer. The effective test Machnumber was obtained fram contour chart8 similar to thosg presented in-figure 5 from the relationship The vaxiation of mean test Reynolds number with Mach number is shown in figure 6, The boundaries in the figure indicate the range in Reynolds number caused by variatiorm in test conditions during the course of the investigation. Force and moment data, effective downmeh anglee, and the ratio of dynamic preeeure at 25 percent of the mean aerodynamic chords of the free-floating tails to flree-tream dynamic pressure were obtained for UCA RM L9G27 - 5 the model confi-guratione,tested through a Mach number range'of 0.60 to 1.18 and.an angle-of-ttack range of -2O to 8.O. The end-plate tares on drag were obtained through the test Mach n'lrmber range at zero angle of attack by testing the model canfiguratims without end plahs. .For these tests a gap of about 1/16 inch was maintained between the whg root asd the bump .surface, and a sponge- xipep seal waa faatened to the wlng butt beneath the surface of the hw to prevent leeas (fig. 4(b) ) . he drag end-phte tares were aas-d to be invasiant with angle of attack .and the- tar06 obtained at zero angle of attack were applied to dl drag data. Je" cnrrectione have not been evaluated inaemuch as the boudaq conditions to be satisffed are not rieously deffned. However, inR.nmuch aa the effectiye flow field is lmge cmgared with the apas d chard of the model, these correctiom are believed to be Bmall. The possibility of change In aerodynamic characterTstics of the wing due to twlst -resulting fram bending mder aerodpmmic 10- con- sidered. Fram etatic loading tests and reference 6, it was estimsted that. at the highest Ma& number attahed the effect of twist would be .to cause a farvd Eier-c-ntermovement .of about 2 percent. RO c&rectiane have been applied to the data presented. From meas-nts of 'tail floating angles xlthout a delFzlEtdled, it wss determined that EL tail spacing of 2 inches relative to the wing . chord plane would produce negligible interference effects of reflected shock waves rn the tail floating angle's. Downwaah angles for fhe alone configuration were therefore obtained simultaneously f& the middle, highest, and lowest tail positions in one series of teste and for the t'wo Intermediate positiorm in succeeding rum. (See ' fig. 3. ) For the wi-elage test8 the effective downwash &ngles'at the chord plane extended were determhed by.munting a free-float* tail on the center lfne of .the 'fuselage. The downwash angles presented are increments from the tal floating angle6 without a model in position. It should be noted that.the floating angles measured are actmUy a measure of the angle of zero pitching m&nt about the tail-pivot axis rather than of the angle of zero lift. It'- been esthated that, fathe tail. arrapgement used, a 2O spaswise domwash gradient over the tail win I result in an error of less than 0,.2O in. the resultant floating angle. Total pressures obtained *can the tail amey rake have been corrected for bow-wave loss. The static-gressure values. used in cmputing dynamiypre6.sure ratios were 0btalned.b~use of a static probe . without a model In position. 6 NACA RM ~9~27 A table of ‘the figures presenting the results follows: Figure Wimlonefwce data ....................... 7 Win&melage fmce data .......................8 Effect of--~fngfence (wing-fuselage) ....................9 Effective downwash angles (wing alme) ....... ; ....... 10 Effective dovnwash angles (wing-fuselage). .............ll Downmh gradients ...........................12 DynamZvrassure surveys ... ..-. ................13 summary of wrodynamic characteristics ...............14 Unlees gated, the &iscussion iB based on the amunary curves presented In figure 14.