at altitudes between 30,000 feet and 20,000 feet.

Transcription

1 NACA RESEARCH MEMORANDUM FUGHT MEASUREMENTS OF THE VERTICAL-TAIL LOADS ON THE CONVAIR XF-92A DELTA-WING AIRPLANE By Clinton T. Johnson High-speed Flight Station Edwards, Calif. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON October 27, 1955

2 .D ETACA RM H55H25 NA!TIONAL ADVISORY COMMITTFZ FOR AEEONAUTICS. FLIGHT B S OF THE VERTICAL-TAIL LOADS OB By Clinton T. Johnson SUMMARY The aeroaynamfc loads acting over the vertical tail were determined from steady and maneuvering flight during the investigation of the lateral stability and control characteristics of the Convair XF-92A airplane. The results presented in this paper were obtained from rudder pulses and gradually increasing sideslips over Mach the number range from 0.50 to at altitudes between 30,000 feet and 20,000 feet. The vertical-tail panel bending-moment and normal-force chazacter- - istics are essentially linear with increasing sideslip angle both in rudder-fixed and trfmmed maneuvers. A comparison of the bending-moment and normal-force pazameters derived from rudder-fixed oscillations and the corresponding parameters derived from gradual manuever8 indicates similar trends with Mach numker. The effect of rudder deflection to is reduce the slope of the vertical-tail normal-force-coefficient variation with sideslip angle and to move the lateral location of the center of pressure of the additional air load inboard about 5 percent of the span of the vertical-tail panel. Phe vertical-tail bending-moment and normal-force coefficients resulting from rudder deflections me essentially constant below a Mach number of 0.80 with an apparent tendency for both parameters to increase at the higher Mach numbers tested. As part of the cooperative Air Force-Navy-NACA flight research program, the delta-wfng Convair IIF-ga airplane was utilized for flight investigations at the NACA Htgh-Speed Flight Station at Edwards, Calif. The primary purpose of these flight investigations was to evaluate the handling qualities, lift and drag characteristics, aerodynamic loads

3 and load distribution, control surface loads, and buffeting characteristics. During the test program the flight envelope was extended the to maximum lift and Mach llumber attainable. Results of several of these investigations are reported in references 1 to 4. Vertical-tail loads were measured by strain-gage methods during these flight investigations to provide full-scale flight loads information on a low-aspect-ratio triangular vertical-tail configuration such as used on the XF 92A airplane. This paper presents the results of the measurements of vertical-tail loads during rudder pulses, rudder-fixed oscillations, and gradually increasing sideslips at level-flight lift coefficients at altitudes between 30,000 feet and 20,000 feet up to a Mach number of II SYMBOLS bvt Cbvt span of vertical-tail panel outboard of gage station, in. vertical-tail panel bending-moment coefficient, C bgr variation of vertical-tail panel bending-moment coefficient acb, with rudder deflection, per deg, - hr Lvt vertical-tail panel normal-force coefficient, - s% variation of vertical-tail panel normal-force coefficient with angle of sideslip for zero rudder deflection, per C Nsr variation of vertical-tail panel normal-force coefficient =N, with rudder deflection, per deg, - ar chord at any section, in..

4 . - cvt NACA RM H55H25-3 mean aerodynamic chord of vertical-tail panel, in. - dc Nvt slope of normal-force-coefficient variation with sideslip - de &le for trimmed sideslip, perdeg e! A Lvt lateral location of the center of pressure of the additional air load of the vertical-tail panel, percent bvt pressure altitude, ft vertical-tail panel aerodynamic load (positive load to the left), ~b I M free-stream Mach number %vt P vertical-tail panel bending moment about vertical-tail straingage station (positive conterclockwise when viewed. from the re-), in-lb rolling velocity, r&ans/sec B free-stream m c pressure, lb/ft2 1: yadng velocity, radians/sec svt area of vertical-tail panel outboard of strain-gage station, f t2 t sec time, Y spanwise distance along vertical tail, in. B. indicated sideslip angle, deg 6r rudder position, deg

5 4 NACA RM H55H25 The Convair XF-92A is a semitailless, delta-wing airplane with a 60 leading-edge sweepback of the wing and vertical stabilizer. The wing and vertical tail have a streamwise thickness ratio of 6.5 percent, The elevons and rudder are =-span, constant-chord surfaces small, with unshielded horn balances near the tips. Control surfaces are actuated by a 100-percent hydraulically boosted system. The airplane has no dive brakes and no leading- or trailing-edge flaps or slats. A three-view drawing of the airplane shown is in figure 1 and photographs are shown in figure 2. Table I lists the physical characteristics of the airplane. INSTRUMENTATION AND ACCURACY The Convair XF-9211 airplane was equipped with standard NACA recording instruments for recording the following quantities pertinent to this investigation: Airspeed Altitude Normal and transverse acceleration Rolling angular velocity and acceleration Yawing angular velocity and acceleration Control positions Angle of attack and angle of sideslip A multichannel oscillograph was used for recording strain-gage outputs and a common timer was used to correlate all instruments. Strain gages were installed on the vertical tail spars at the vertical tail root (approximately 4 inches outboard of the vertical tail-fuselage Juncture as shown in fig. 1) to measure shear and bending moment. The data presented in this paper have been corrected for the inertia of the vertical tail and are the aerodynamic loads acting over the vertical-tail surface. The accuracy of the measured was loads determined from the results of a static calibration and evaluation of the strain-gage responses in flight. The estimated error in shear and bending moment 5300 pounds is and?8,000 inch-pounds, respectively. Estimated accuracies of other pertinent recorded quantities : are

6 - p, NACA RM H55H25-5 Mach number, M r, radians/sec... "0.02 radians/sec... f0.10 p, deg... H.25 6r, deg... S.20 TESTS The flight tests were conducted in the clean configuration at levelflight lift coefficients. VertTcal-tail loads were measured during abrupt rudder pulses, rudder-fixed oscillations the with aileron held fixed, and gradually increasing sfdeslips using ailerons to hold constant heading over the Mach number range from 0.50 to 0.87 at altitudes from 30,000 feet to 20,000 feet. Reyno ds number, b sed on the wing mean aerodynamic chord, varied between 25 x 102 and 50 x 102 for thfs series of tests. The center of gravity varied between 2'7.1 and 28.2 percent of the wing mean aerodynamic chord. RESULTS AND DISCUSSION Time histories of representative rudder pulses at several Mach numbers are presented in figure 3 showing the rudder input, the resulting vertical-tail loads, and airplane motions.. The initial portion of the maneuvers shows the rudder deflection and the corresponding change in vertical-tail bendfng-moment and normal-force coefficients before the airplane responds to the control input. This portion of the maneuver is indicated by the solfd lines in 3 figure occwing near t = 1.0 second. Since rudder deflection, vertical-tail bending-moment, and normal-force coefficient were the only variables during this portion of the maneuver, it was possible to determfne the vertical-tail-load parameters cbsr and C N ~ ~ It. may be noted that small changes in sideslip angle (less than 0.lo) did occur during the time the mrudder s being deflected. However, the error in the values of CN and Cb, caused by a change Sr 6, of 0.1O in sidesllp angle, was estimated to be less than 20 percent based on the values of normal-force-curve slope and the center of pressure of the additional air load ascertained from this investigation. The Mach number variation of the vertical-tail-load parameters and C determined from rudder pulses is shown in figure 4. '%r The parameters C and C are relatively constant below E Mach % Nsr '

7 number of 0.80 at levels of about per degree and per degree, respectively. At the higher Mach numbers a slight increase in both parameters is apparent. The soiid lines in the latter portion of the maneuvers beginning near t = 2.0 seconds of figure 3 show the airplane oscillations after the rudder has been returned to neutral. From this portion of the maneuver the vertical-tail normal-force coefficient C was plotted with %t respect to sideslip angle p and appeared to have a linear variation with sideslip over the range of sideslip angles investigated. Therefore, slopes of these data were taken to determine the parameter CNB. (Normal- force increments caused by rolling and yawing velocities were evaluated and were found to be negligible). Typical plots used to determine and the variation of CN with Mach number are shown in figure 5. The P vertical-tail-load parameter is constant at a level near CNP per degree to a Mach number of 0.70, then increases gradually to a value near per degree at M = The center of presmre of the additional air load CPA for the rudder-fixed oscillations was determined by taking slopes of the variation of bending-moment coefficient with Cm. Typical plots used to determine cpa and the variation of cpa with Mach number me shown in figure 6. The lateral location of the center of pressure of the additional air load is located at approximately 43 percent of the span of the vertical-tail panel over the Machnumber range from 0.50 to The vertical-tail loads measured during gradually increasing sideslips over the Machnumber range from 0.50 to 0.85 me shown in figures 7 to 9. It may be noted that sideslips were performed using sufficient aileron to hold 8 constant heading. Aileron angles varied from approximately bo at low speeds to 2O at high speeds. The parameters determined from these maneuvers &re compared with the parameters obtained from the rudder-undeflected maneuvers to illustrate the effect of rudder deflection on the vertical-tail loads. Figure 7 shows the variation of the vertical-tail normal-force coefficient C N with ~ sideslip angle and the corresponding rudder required to sideslip for several maneuvers over the Machnumber range. The data of figure 7 are shown in figure 8 as the miation with Mach number of the slopes of the vertical-tail normal-force coefficient

8 resulting from trimmed sideslips - and the ratio of rudder deflecdp tion to sideslip angle The parameter - dcnv-t for trimmed sideslip a dp has a value of approximately per degree to a Mach rider of 0.75, then increases gradually to per degree st M = 0.@. The curve of C determined from rudder-fixed maneurns (fig. 5) % 1s also shown on figure 8. The difference in level between the curves of C and 3 dcet illustrates the change in normal-force-curve slqe % dp ' attributable to rudder deflection and is relatively constant over the Mach number range. The reduction in rudder-fixed % is approximately 20 percent. B It is interesting to note that the variation of -%t with Piach I number can be derived using from figure 4, 5 from figure 8, dp from figure 5, since - = Cwsr x as a, + CNp. This method and C % dp was used to calculate the variation of vertical-tail normal-force coefficient with sideslis angle and agreed very closely Uith the measure data and slopes of figure 7. The variation of bend--moment coefficient cbv% force coefficient CN* db -r with normal- and the resultant cpa for the trimmed side- slips is shown in figure 9. The center of pressure of the additional air load for trhmed the sideslips is located at approximately 40 percent b,t over the Mach number range from 0.50 to A comparison of the centers of pressure of the additional air load determined from rudder-fixed and trimmed sideslip maneuvers indicates that rudder deflection moves the cpa inboard approximately 5 percent of bvt over the Mach number range tested. CONCLUSIONS Flight measurements of the vertical-tailoads on the Convair XF-92.A a rplane over the Mach number range from 0.50 to 0.87 during

9 8 - NACA RM H59H25 rudder pulses and gradually increasing sideslips have indicated the following conclusions: 1. The vertical-tail panel characteristics are essentially linear throughout the angle of sideslip and Mach number range tested. 2. The vertical-tail load parameters derived from the rudder-fixed oscillations and steady sideslip.maneuvers display similar trends with Mach number, with differences in level indicating the effect of rudder deflection on the vertical-tail loads. The predomfnant effect of rudder deflection on the vertical-tail loads is to reduce the normal-force curve approximately 20 percent and to move the center of pressure C% of the additional air load CPA inboard approximately 5 percent of the span of the vertical-tail panel. 3. The vertical-tail bending and normal-force coefficients resulting from rudder deflections, C, and C are essentially constant below 6, N6r a Mach number of 0.80 with both parameters indicating a tendency to increase at the higher Mach numbers tested. High-speed Flight Station, National Advisory Committee for Aeronautics, Edwards, Calif., August 15, 1955.

10 NACA RM H55H25 9 t 1. Sisk, Thomas R., and Muhleman. Duane 0. : hteral Stability and Control Characteristics of ConvaFr the XF-92A Delta-Wing Airplane 8s Measured in Flight. NACA RM E55AJ-7, Sisk, Thomas R., and Muhleman Duane 0. : Longitudinal Stability Characteristics in Maneuvering Flight of the Convair XF-92A Delta- Wing Airplane Including the Effects of Wing Fences. RM NACA E9J27, Kuhl, Albert E., and Johnson, Clinton T.: Flight Measurements of Wing Loads on the Convair XF-W Delta-Wing Airplane. NAa Fibs H55D12, Johnson, Clinton T., and Kuhl, Albert E.: Flight Measurements of Elevon Hinge Moments on the XF-92A Delta-Wing Airplane. NACA RM H54J25a, 1955.

11 10 NACA RM H55H25 win&: b2) m. ft Area. spit Airfoil aectfon... K4C4 6)(&)-rnE.> maerodynaniccbord. fi w Aspectratio Rootchord. ft Tip chord Taperratio... 0 Susepbsck (leading we), dag Incidence... deg 0 ni.tchordplarac), dsg... 0 BLarona: Area (total, both, 2.eQ1yBTd of hinge Uno). sq it ~oru balance arm (tow, both, elevma forvard of him line), sq it 1.4 Span (one elevon), ft Chord (merdard of h m line, mtant except at tip), it... 5.W Idovmnt, de& Elevator 1 up Dovn... 5 Aileron, toid 10 operstion Hydraulic Area, above fuaelags canter line, aq it... S.26.5 )-res,erpoad, aqft Helght. above fuselage center line. ft Heigbt,expossd,ft Airfoil section.... UCA 4(&)-"6.5 man aeroqynamic chord (area nhme fuselage canter lime). in 167.1: MOM aeroqnanxlc chord of exposed area, ln l29.40 Aspect ratio Of area above fuaersgc center lim bpct ratio Of exposed area... 1.w Rwt chord at fuaelage center lins. in... p5l.l5 Rwtchoreofexgoeed.arss,in.... Tipc hord... % TaJ)er.tio... 0 Smpbaek (landing a*), dag Vertiosl tail: Vertical tsil we1 outboard of ge# station: Area, aq Span,in liesnueroqynamiccbrd.in... a.7 FuaeLaga station of la- edg. of w u aerodpwdc chord. in m Tail length to -1 quarter chord at gross-might center of gravity, in hpctratio Root chord, in Rudder : Ares, aq it... l>.z Span, ft... 9.Z Tranl, deg... f6.5 operation... Byluaullc Fuselage: Length, ft ~(ariuumdiamter, ft Power plant: Engine... Rating: "inon XSA-23 with stterburner Statlc thruat at ass level. lb... 5,- Static -at at w lwei with afterburnu, lb... 7, YXl Weight: Groasueight($60&alfuel), lb w R&yveight. lb... U, 8oti Center-of-&ravlty locations: m-osu we & (* gal mol), parcent H.A.C myuelght, gcrcentii.a.c.... e.2 M i a characteriatica: w t of irrsrtia in ynw, (avey.valua thrawh 00 to msle of attack at average 50sa nignt). slug-it... 38, 400 m n t of imrtia in roll. (avorase value thrcm& 00 to an6le of attdck at avorags groaa weight), a1ug-d... 6, 000 hclimtiotl Of grincim Of inartis, (est-td), deg

12 . - m Figure 1.- A three-view drawing of the XF-92A airplane. All dimensions are in inches.

13 12 (a) Left side view. L (b) Three quarter rear view. (c) Overhead front new. Figure 2.- Photographs of XF-92A research airplane.

14 NACA R4 II55H25 Right 2 p,mdians&c 0 2 Time,t, sec (a) M = 0.52; hp = 23,000 feet. Figure 3.- Time histories of airplane motions and vertical tail loads resulting frm typical rudder pulse maneuvers at eeveral Mach numbers.

15 14.OS.04 %t CNvt.I 0 -.I Right 2 p,rodlanskc;ec 0 2 Right.2.2 Right 4 f4 deg 0 Time, t, sec (b) M = 0.71; % = 31,OOO feet. Figure 3. - Continued. Y

16 NACA RM H %t "08 L Time, t, sec (c) M = 0.87; hp = 30,000 feet. Figure 3.- Concluded.

17 16 NACA FM H55H C, per deg % Figure 4.- Variation with Mach number of vertical-tail bending-moment and normal-force coefficients caused by rudder deflection. M

18 3 NACA FM E55EE Right. 1 Figure 5.- Variation of vertical-tail normal-force coefficient with sideslip angle, and the normal-force curve slope variation with Mach nuber during rudder-fixed oscillations. M

19 18 U A OM=Q87 f7 M=0.71 o M=0.52 c %t CPA, percent b,t Figure 6.- Variation of vertical-tail bending-moment coefficient with. normal-force coefficient and the variation with Mach number of the center of pressure of the additional air load during rudder-fixed oscillations.

20 NACA RM H55E Left. (a) M = 0.52; kp = 29,000 feet. Figure 7.- Variation of rudder deflection and vertical-tail normal-force coefficient with sideslip angle from several representative trfmmed sideslip maneuvers.

21 X) P, deg Right (b) M = 0.72; % = 30,OOO feet. Figure 7. - Continued. I

22 NACA RM H55H I. (c) M = 0.85; kp = 22,400 feet. Figure 7.- Concluded.

23 22 NACA Rt4 H55H25 per deg Figure 8.- Variation with Mach number of the rudder required to sideslip and the vertical-tail normal-force cue slope from trimmed sideslip maneuvers showing the effect of rudder deflection. M

24 Rudder undeflected rc Trimmed sideslim Figure 9. - Variation of vertical-tail bending-moment coefficient w it;h normal-force coefficient during trimmed sideslips, and the variation with Machnumber of the center of pressure of the additional air load showing the effect of rudder deflectfon. M

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