1、NATIONALADVISORY COMMITTEEFOR AERONAUTICSTECHNICAL NOTE 3314A TECQUE UTIIJZING ROCKET-PROPELLEDTEST VEHICLES FOR THE MEASUREMENT OF THE DAMPING IN ROLLOF STING-MOUNTED MODELS ANDDELTA AND UNSWEPTBy William M. Bland, Jr.,SOME INITIAL RESULTS FORTAPERED WINGSand Carl A. SandahlLangley Aeronautical Lab
2、oratoryLangley Field, Va.WashingtonMay 1955Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1s TECHLIBRARYKAFB,NM :IWTIOIWL ADVISORY COMMITTEE FOR AERONAUTS IllllllllullllllullllllllllI00LL032 jTECHNICAL NOTE 3314A TECHNIQUE UTILIZING ROCKET-PROPEILED
3、TEST V12W2LES FOR TEE MEA OF TEE DAMPING IN ROLLOF STING-MOUNTEDMODELS AND SOME INITIAL RESULTS FORDELTA AND UNSWEPT TKPERED lCD12S1I however, the agreementa71 improved with increasingMach nuniber. Ihcreased section thicknessdecreased the damping in roll of the delta wings throughout the Machnuaiber
4、range investigated.“INTRODUC!KIONThe Langley Pilotless Aircraft Research Division is utilizing twoexperimental techniques emplng rocket-propelled test vehicles forthe determination of the -ing-in-roll derivative at high stisonic,transonic, and supersonic speeds at relatively large Reynolds nwibers.O
5、ne technique which is used for determining the demping in roll.of wing-fuselage cozibinationsis described in reference 1. The other techniquewhich is used for determining the damping in roll of wings alone and ofwing-fuselage conitxinationsis described herein. The Reynolds numbersobtained with the u
6、se of this technique, although somewhat lower thanthose obtained with the technique of reference 1, are still fairly high(1x 106 to 3 x 106).6%upersedes re6ently declassified lUlCARML50D24, 1950,sProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA
7、 TN 3314Also presented herein are scme initial results obtainedby thepresent technique for a series of configurationshaving wings of aspectratio 4. The configurationsinvestigated included a delta-wing-fuselage conibinationhaving a wing made frgm a flat plate wi.thbeveledleading and trailing edges, t
8、wo delta wings-having 45 of leading-edgesweep - one with a 4-percent-thick symmetricaldouble-wedge airfoil sec-tion and the other with a g-percent-thicksymmetrical dcnible-wedgeairfoil section, and an unswept tapered wing hav, 0.5 taper ratiowith a 4.6-percent-thick symmetrical double-wedge airfoil
9、section.clPC2pbzLsbPvMRczYSYMBOLSdCzdamping-in-ro derivative,nd pbELrolling-mment coefficient, qsbwing-tip helix angle,rolling moment, ft-lbradisnsdynamic pressure, lb/sq ftwing area, sq ftwing span, ftrolling velocity, radiarm/secflight-path velocity, ft/secMach numberReynolds numiberwing chord, ft
10、based on wingwing mean aerodynamic chord,lateral coordinatemean aerodynamicz b/2J730 c2dy, ftchordProvided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 33149t thickness, ftw A aspect ratio obtained by extending wing leading and trailtigedges to mod
11、el center lineA leading-edge sweep angle, degh taper ratioMETHODThe general arrangement of the test vehicle is illustrated in fig-ures 1 and 2. The wing under investigationwas attached to a torsionspring balce arranged to form a sting mount in the nose of the testvehicle. In flight the entire test v
12、ehicle was forced to rollby thestabilizing fins, each of which was set at an angle of ticidence. Arocket motor accelerated the test vehicle to the maximum Mach number,after which the test vehicle decelerated through the test Mach nuniberrange. Time histories of the rolling moment generated by the te
13、st wing,the flight-path velocity, and the rolling velocity were obtained. Thesedata, in conjunction with atmospheric data obtained by radiosonde meas-urements, permitted the evaluation of the damping-in-roll derivative%P as a function of Mach nuniber.a71A ssmple flight path illustrattig the useful.r
14、ange of a flight andsome tical conditions is shown h figure 3. Typical time histories.of some of the measured qusmtities are shown h figure 4.A photograph of a test vehicle mounted on the zero-length launcheris shown in figure 5.INSTRUMENTA!TIONThe torsion spring balsnce shown in figure 6 consisted
15、of a shsftwhich transmitted the rolling moment generated by the test wing to ahelical torsion spring which permitted angular movement relative to thetest vehicle proportional to the rolling moment. The angular movementsof the shsft were transmitted to a condenser-typepickup which was usedin conjunct
16、ion with standard NACA telemetry.The rolling velocity was obtainedby the method of reference 2.except that the telemeter and telemeter antenna performed the functionsof the spinsonde described in the reference. The telemeter antennaa71Provided by IHSNot for ResaleNo reproduction or networking permit
17、ted without license from IHS-,-,-4 NACA TN 3314consisted of two rods which were inserted in the trailing edges of two a71diametrically opposed driving fins as shown in figure 1. This antennaarrangement produced the plane polarized radim-signal required for themethod of reference 2. The ground record
18、ing equipmentwas the same as *that described in reference 2.The flight-path velocity was measured by a Doppler radar veloctieter.The altitude, which was obtainedby integratingthe velocity-time curve,was correlated with radiosonde measurements ofalong the flight path made at the time of eachTEST CONF
19、IGURATIONSatmospheric conditionstest flight.The configurationstested, all of which had an aspect ratio of 4.00,were (1) a delta-wing-fuselage combination employing an airfoil sectionhaving fla,tsides and symmetricallybeveled leading and trailing edges(f. 7), (2) a delta- having45 of leading-edge swe
20、ep witha-k-percent-thick symmetrical double-wedge airfoil section, (3) a delta wing having 45 of leading-edge sweep with a g-percent-thick symmetricaldouble-wedge airfoil section, and (4) a wing having a taper ratio of 0.5with an unswept 50-percent-chordline and a 4.6-percent-thicksymmetricaldole-we
21、dge airfoil section.The gecxnetriccharacteristicsof the con-figurations tested are summarized in table I. Photographs of the testconfigurations are shown in figure 8. The wing surfaceswere carefullyground and polished after being machined from steel plate. The distancefrom the trailing edge of the r
22、oot chord of the wings to the nose of the a71test vehicle was-held constant as showz.in figure 1:ACCURACYThe maximum possible systematic errors in the valuessented herein due to the ltiitations of the measuring andsystems are estimated to be within the following limits:.of cl p pre-recording Delta w
23、ings Unswept tapered wing I ,M Error in Cz M Error in CzP P1.7 *().008 1.7 “- +0.0151.4 *.013 1.2 *.030k.033 1.0 * .041:? *.053 .7 * .100Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA !IIT3314 5a71 The variation of these possible errors is due
24、to errors of constant_itude cuded ti some of the measured values; therefore, whereverw the measured rolling moment or roll velocity, or both, decreases,the respective errors become a larger part of the measured values andincrease the possible error.An error in the determination of Czp may exist beca
25、use of thenecessity of neglecting the tare rolling moment; that is, the rollingpbmoment which might exist at = O because of inaccuracies in model con.struction. However, the results obtained for mdels with only nominaldifferences presented herein agree well within the aforementioned limits.Any contr
26、ibutions to the possible errorby the drag of the testconfigurations and temperature variations on the spring balance are neg-ligible when compared with the errors previously tabulated.The measured rolling moment included a moment equal to the productof the moment of inertia of the test assembly (win
27、gs and contributingparts of the torsion spring balance) and the instantaneous accelerationin roll. In the present investigation the inertia rolling moment pro-duced a maximum error in Cz of -0.002; therefore, the data are pre-Psented without correction for this error.RESULTS AJiDDISCUSSIONThe result
28、s for all the wings investigated are presented in fig.ure 9 as curves of rolling-moment coefficient CZ, wing-tip helix singlepbj ud rig-in-roll derivative C% as functions of Mach rnmit)er.The results presented in figures 9(b) however, the agreement improveswith increasingMach nuniberand decreasing t
29、lgicknessratio. In thestisonic speed range the results agree within experimental accuracywith values from reference 5 to which appronat corrections for theeffects of compressibilityhave been applied by utilizing the Gla,uert-Prandtl transformationas described in reference 6.Unswept %aperedWingFigure
30、 n(b) compes the variation of CZ with Mach nuniberPobtained for the unswept tapered wing with theoretical results and shows . s_that demping in roll.is maintained throughout the Mach nuniberrangeinvestigated. At subsonic speeds the agreem the agreement improved withincreasingMach mxiber. Decreasing
31、the thickness ratio of the deltawings improved the agreement with the linearized theory.Langley Aeronautical Laboratory,National Advisory Committee for Aeronautics,Lazey Field, Vs., May 2, 1950.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACA T
32、N 3314REFERENCES ?.1. Edmondson, James L., snd Ssnders, E. Claude, Jr.-: A Free-Flight.Technique for Measuring Damping in Roll by Use of Rocket-PoweredModels and Some Initial Results for Rectangular Wings. NACARM L9101, 1949. 2. Harris, Orville R.: Determination of the Rate of Roll of a71PilotlessAi
33、rcraft Research Models by Means of Polmized Radio Waves. NACATN2023, 1950. . 3. Lomsx, Harvard, snd Heaslet, Max. A.: Dsmping-in-RollCalculationsfor Slender Swept-RackWings and Slender Wing-IkxlyCombinations.NACATN 1950, 1949.4. PiA = 45i=a Delta;b A = 45a Delta;b A n k513c/2 line Wlenepttaper ratio
34、, O.51BodywithirithmtaithcultJithoutWngihicknemratio0.040.Ogo,046AkeoilsectionEeveledplate9gmmatricaldoubleV*3ymetricaldoublewedge3ymmtrlcaldoublesing erea( aqsrt)0.IB3.I.$!3.I.8e, leakpect ratio,Wng span,A (L)+4.00 I 9674.00 .%7Meful-chord,(L)0.288.2a9.289.225Provided by IHSNot for ResaleNo reprodu
35、ction or networking permitted without license from IHS-,-,-920.0/-+.-;.,I5D I“L =.4,-h=i-.30d45”-I.-.:- :.- _=-. ”-,-,General arrangement of a model with all dimensions in inches.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NAC.Am 3314 l-lGu.a71.r
36、-=- -.:- .,- . . . ,.-”. . -jiFigure 2.- ry-picaltest vehicle-=s=L-61603.2with test wing attached.1Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I-2- mm m 3314/ 10 secaccelerating flightcoctsting flightaL-41/ v “ “ “-”:/ -.Figure 3 Sample flight pa
37、th with perfoce figures.8Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 3314 130 I 2 3 4 5 6 7 8Time after launching, secondsFigure 4.- Variation with time of flight-path velocity, Mach nuuiber,rolling velocity, and rolling moment for ccmfig
38、uration 2(b).Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 NACATN 3314v, . . . -., .-” “-k- . .“ -. -”-r- - .-1-.- - -e. &*-. - - . . _ .:.:”,-,. 1AF , - ,=,%&=+.,.Figure 5.- Test vehicle1*-on zero-length.w“-”-L-61812.1 a71launcher.Provided by I
39、HSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NAC!Am 3314 15Front support & Torsion shaftH#icaql tors-bnFigure 6.- Nose cone interior showing spring balance aud capacitancepickup.Provided by IHSNot for ResaleNo reproduction or networking permitted without licen
40、se from IHS-,-,-NACA TN 3314145L 10.35”t,#D.+4-15,9” 0857”.-,.“ -. .Figure 7.- Configuration 1. Fuselage ks a circular cross section throughwhich the wing passes diametrically. wing was made of 0.146 steelplate with beveled leading and trailing edges.Provided by IHSNot for ResaleNo reproduction or n
41、etworking permitted without license from IHS-,-,-3s.*.NACA TN 3314 17(a) Configuration 1.Figure 8.- Configurations tested.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 3314.(b) Configurations 2 snd 3.Figure 8.- Continued.Provided by IHSNot
42、for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACATN 3314(c) Configuration 4.Figure 8.- Concluded.19.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-20 mcll m 3314.016 - / - G#08o 5 7 .8 .9 l-o 1.1 I-2 I-3 1.4 I*5 I*6
43、 J-7 -M.06.04 g.02I I t I I 1 i I I 1 I I I I I I t I I Iu 1 , , I I , i , , I , , J5 7 .8._.,9 I.O 1.1 1,2 1.3 1.4 1.5 !.6 1.7M-4-.3-.1n.M/ 1 ./ - -= ./ - _- 6 7 .8 .9 1,0(a) Configuration 1, 45Figure 9.-1.1 1,2 1,3 14 !,5 1,6 1,7Mdelta-wing smd fuselage combination.Experimental results. .Provided
44、by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 3314 21Modela . . b.024.016%.0080 /, - - - - N- . . . , _ _ -. ._ _-.6 .7 .8 .9 1.0 1.1 1,2 .3 1,4 ,5 166 1.7M.06a71.04- = .- - . -.-.- -, _ ._. _ ,_bh.020 ,6 .7 .8 .9 1.0 1.1 1.2 1.3 1 1.5 I-6 1.7M-.3C$
45、2-. Io- - / “ “ - -.-,1 . T+-. - _- - - - .=E=- .6 .7(b).8 .9 1,0 1,1 ,2 1.3 .4 1,5 1.6 I*7MConfiguration 2, delta wing, A . 45, */c = 0.040.Figure 9.- Continued.Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-mcfl m 3314Modelo- -. b.024.016 - - - .
46、008 .- - - L -, J - - “- = = = = = =- f0 ,6 .7 .8 .91.0 1.1 1.2 1.3 1.4 1.5M.08 1 - - _.06 -1.04- -“- - .02o6 .7 .8 .9LO 1.1 1.2 1.3 1A 1.5M-.3-, 2%p-. to /u-=Y9=-S .7 .8 .9 1.0 LI 1.2 1.3 1A .5M(c) Configuration 3, delta wing, A = 45, t/c = 0.090.Figure 9.- Continued. .-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA TN 3314 23.6 .7 .8 ,9 1.0 1. 1.2 1.3 1.4 1,5 1.6 1.7M.06.04 -b%?020 ,6 .7 .8 ,9 1.0 1.1 1.2 1.3 1.4 1.5 .6 1.7M-.6-.5-A-.2-, Io.6 .7 .8(d) Configumtion.9 1.0 1.1 1.2 1.3 1,4 1.5 1.6 .7M4, unswept tapered w
