ARR No. L5HII y NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED October 19^5 as Advance Restricted Report L5HII COMPARISON BETWEEN CALCULATED AND ME/VSUKED LOADS ON WING AND HORIZONTAL TAIL IN PULL-UP MANEUVERS By Cloyce E. Matheny Langley Memorial Aeronautical Laboratory Langley Field, Va. WASHINGTON NACJA W,\RTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- nicajly edited. All have been reproduced without change in order to expedite general distribution. 193 Digitized by tine Internet Arcliive in 2011 witli funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/comparisonbetweeOOIang ;j3>^^ ^ :?r.> 3?r. NACA ARR :]o. L5H11 NATIONAL ADVISORY COI.IMITTEE FOR AERONAUTICS ADVANCE RESTRICTED REPORT C0KPARI30N BETVi'EEN CALCULATED AND MEASURED LOADS ON WING AND HORIZONTAL TAIL IN PULL-UP MANEUVERS By Cloyce E. Matheny SUMM'IRY Comparisons have been made of measured and calculated loads on the wing and the horL-^ontal tail in pull-up rr-aneuvers for six airplanes ranging iii weight from ix,700 to 48,000 pounds. The calculated loads were based on the control motions r.easured in flight. The aerodynamic char- acteristics of the airplanes required for the calculations were either obtained directly from wind-tunnel data or computed , Good agreement was obtained betv;een calculated and measured loads for a specified elevator deflection 7/hen reliable wind-tunnel data were available and when the airplene maneuvers -vere consistent with the assumptions. The fact that only fair agreement was obtained in some of the cases was attributed either to poor quantitative knowledge of the aerodynain.ic parameters or to the viola- tion of the assumptions on v;hich the method is based. INTRODUCTION ■ During the past few y/ears rf.uch v.'ork has oeen done in an attempt to relate tail loads more closely to the aerodynamic and geometric characteristics as well as to the functional requirements of the airplane . In various reports that have been written on this subject either of tw^o approaches has been used: namely, (1) to proceed from a specified control motion to the determina- tion of the wing and tail loads, as in reference 1; or (2) to proceed from a specified wing-load variation to the determination of the tall load and elevator motions, as In reference 2. Both methods depend on a solution of the equations of motion for a rigid body and consequently require a knowledge of the aerodynamic and geometric NACA ARR No. L5H11 characteristics of the c^irpltvne. The loads computed are the rejiiltant air loads that act over the horizontal surface-,; thei-efore the eolntions chtalned do not irdioate possihlc adverse chordv/ise or spanv;is3 distributions ox' the buffeting tall- load increment- Rocently a :aethod based on the determination cf the wing and tail loads for a soeclfied control notion has been recommended as a part of the' airplane load design requirements for the Army (references J and 1+) . Since the application of this method reqidres considerable time, it seems desirable to deterxaine the agreement that can be expected between measured and calcialatod results. The object of the present report is to give results of comparisons between measured and calculated wing and tail loads in pull-up maneuvers for six airplanes ranging in v;eight from I,., 700 to i+SjCOO pounds. The flight data presented heroin are tvpical and are taken frcm unpublished results measir.red in flight during the past fire Tears. SyiVIBOLS W airplane v/eight , pounds g acceleration of gravity, feet per second'^ m airplane mass, slugs ('jV/g) S gross wing area including area within fuselage, sq^lare feet S-j; gross horizontal-tail area including area inter- cepted by fuselage, square feet b wing span, feet b^l; tail span, feet ky radius of gyration about pitching axis, feet Iv moment of inertia about pitching axis, slug-feef^ Xx. length from center of gravity of airplane to uerodynLOiiic center cf tall (negative for con- ventional airplanes), feet NACA ARR No. L5!!ll 5 o air dens-.ty ratio ( p/p . ) airspeed, feet per second eqviivalent airspeed, rniles per houi' ( 1 M \.l. 14.07 M Mach n-ainber p mass density of air, slugs per cubic foot q dTnami c pressure, pounds per sq.uare foot 1 ?"'■ cL t fc . /.-A. T] tail efficiency factor /q^/q^ L lift, pounds Cl lift coefficient (L/q3) Cjj,^ pi telling -moment coefficient of airplane v.'ithout horizontal tail [Moment x qS2 a wing angle of attack, radians a^. equivalent tail angle of attack, radians 5 elevator angle, radians /de \ e dovmwash angle at tail, radians I — a) Vda ) K empirical constant denoting ratio of damping moment of complete airplane to dejiiping mom'jnt of tail alcne n airplane load factor K] ', K2'' -'-'^-T ' nondimensional constants occurring in "basic differential equation • • • The notations a and a denote single and double differ- entiations with resuect to time. k FACA ARR No .5K11 Subscripts: t tail sea- level conditions iv'i XL, ^ nu iJ Although, as previously stated, tliere are a number of methods available for computing the wing and tall loads for any elevator notion, the method used herein for all the computations is that described in reference 1. This method is similar, as far as basic assimiptions are con- cerned, to that of reference I; but differs in small details such as t^rpe of axes used and computational pro- cedures employed. The basic ass'i-imptions underlying the m e tho d are tha t • {!) The change in load factor in a pull-up or pull- out, as a result of attitude change, is small with respect to that due to change in angle of attack (2) The aerodjTiamic quantities are linear functions of angle of attack (3) The speed is constant during the maneuver (i|.) The effects of flexibility are neglected With these assumptions the differential equation of motion for a unit elevator deflection becomes *a + Kt'ci + Kp' Aa = K2;' A5(1) ■J- C T (1) where Kn ' 5 ^2' , and K2; ' are functions of the aerody- namic and geometric characteristics, With the unit solu- tion of equation (1) known, Aa and a. are evaluated for any control motion by applying Duhamel's integral theorem. The Increment in load factor An is related to Aa thJT'Ough the equation dC^ Aa q d a w/s (2) NACA APR Ho „ L5H11 5 The increment ^ In equivalent tail angle of attack is related to Ac and a through the equation ^Cl P S xA . x^^,/^e , l\ d ta^^ = Aa(l - and finally the tail load follows fr-on equation (5) as BASIC DATA FOR CALGIILATIOIIS Flight data.- The flight data used in the calcula- tions" are given in figure 1. This figure shovirs the time variation of airspeed and elevator position measured during either pull-uns or cive pull-outs made -t/ith the SB2C-1, PBf.'i"5./ P--lj.Ox[, XP-51, BT-9B, and , B-P.i+D airplanes . The wing and tail loads corresponding to these control motions and airspeeds are included in the figures giving the comparisons tet'7een calculated and measured values. Aerodynamic para m eters . - The aerodynamic parameters required are j slope of airplane lift curve. This quantity v/as obtained, whenever possible, from wind-tunnel tests of eitner the complete airplane or a 0.0-r m.odel. For the PBLI-J seaplane, — — was esti- iCL dCi l.a+ da mated from tects of a mod.el of a similar sea- plane . slope of tail-plaiie lift curve. This quantity was obtained^ v.lienever possible, from wind-tunnel tests of the isolated tail or from tail-on tests made with different stabilizer settings. When such, data were not available from tunnel tests, they were obtained from reference 5- NACA ARR Ko. L5H11 — rate of change of downwash at tail with angle of '^^ attack. This factor was determined, v/henever possible, from results of downwash surveys beliind a particular model or from moment differ- ences between tail-on and tail- off wind-tunnel tosts. When experimental results v/ere not available, this factor was computed from the results given in refersnce 6. Tj tail efficiency factor. When possible, this fac- tor was obtained from total-head surveys in the region of the tail. .Vhen such surveys were not available, the method suggested in reference 1 was used to detennine this quantity. K empirical damping factor, ratio of damping m.oment of complete airplane to that of tail. In the calculations this value was taken either as ].,1 or 1.25, depending upon the airplane configu- ration. "da"' elevator effectiveness. Trds quantity was obtained from reference 5 whenever specific wind-tunnel tests were not available for its determination. slope of airplane moment-coefficient curve (minus tail). This q.uantity vfas determined from wind- tunnel tests of either a model or the airplane. The values obtained from the tunnel were adjusted for the particular center-of -gravity position of the flight tests. ?or the dC™ ?.Br/i.-3 seaplane, — ~ was estimated from tests da of a model of a similar seaplane. rate of change of tail moment coefficient with ^^^ elevator deflection for isolated tail. Except in the case of the BT-9B airplane,, this quantity was computed from results similar to those given in reference 7« The aerodynamic parameters for all airplanes under consideration are comoiled in table I. Low-speed v/ind- tunnel data were used in all the foregoing parajiieters "^Cxrit; NACA hP.R Mo. L5H11 7 except for a few values on the X?-51 airplane, which were taken from v/ind- tunnel data at the flight Mach number. The remaindei' of the parameters for this airplane v/ere corrscted for tne effects of Ivlach nu:n"oer hy the Prandtl- Ciiauert factor — No corrections were made to the (l - m2 lov.'-speed values for the other airplanes. In order to chec]^ the validity of these calculations, an individual case was calculated v/hereby the effects of compressibility v/ere taken into account for a dive by the 3B2C-1 airplane at a Mach number of O.bl. Results frora these calculations showed that at this Mach number the loads calculated usinj parameters corrected for compressi- bility effects were not apprecifably different from the loads calculated using low-speed values of the parameters. Physical and geoi iet ric characteris tics.- The physical and geometric cho.rsctei-i sties "of the airplanes were deter- mined principally from manufacturers ' data and are presented in-" detail in table II. RESULTS- The increments in accalerabion and tail loads com- puted from the basic data given in figure 1 and in tables I ana II are shown in figures 2 to 10. In these figures the dashed lines represent the calculated values and the full lines, the measured values. The measured tail loads v/ere obtained by use of pressure distributions, electrical strain gages, beam deflections, or dynamometers Table III summarizes the tail load conditions represented in the various figures and gives the estimated accuracy of the measvirements . In these comparisons (figs. 2 to 10) the tail loads given as -'measured tail loads" have been converted to air loads: that is, inertia effects have been eliminated when necessary. In each case the comparison is made of the increments in load measured with respect to the loa.ds at the instant the maneuver was considered to have been started . The measured accelerations were obtained with a standard NACA acceleroKeter located near the center of 8 NACA ARR No , L5H11 gravity. The measured accelerations are accurate to about ±0.05g. DISCUSSIOK The comparisons given in figures 2 to L). for the S320-1 airplane indicate good agreeinent between the calcu- lated and measured tail loads for the three typical dive pull-cuts chosen. The measured data were obtained at Maoh nu:nbers below the critical value for this airplane, 0.67? and in relatively quick pull-upe . Such conditions favor the assumptions on v/hich the calculations were based: naraely, linear variation cf aerod^rnamic quantities vvith angle of attack, and small attitude and speed changes during the maneuver, A great deal of consistent wind- tunnel da^a were also available for this airplane in the fox^m of force tests and v\/ake surveys behind a model. Although ohe flight conditions shov/n in figures 2 to I|. are not the critical ones for which calculations would ordi- narily be made, the fact that the calculated and measured tail loads per g are approximately the sanie indicates that the method could be used to predict loads with good accu- racy for condi'cicns other than thoso tested. The cciniDarison shown in figure 5 ^'^^ a pull-up with the PH/'!-3 seaplane sho'ws good agreement in the acceleration increments obtained. This calculation represents one for which a minimum of wind-tunnel data was available. The pull-up was made from a shallow dive and in such a \;ay that both small attitude and velocity change resulted. Although no tail loads were measured in flight on this seaplane, the calculated tail loads are thougnt to be of interest . The results shown in figure 6 for a dive pull-out with the P-I4.OK airplane show poor agreement between the acceleration and tail-load increments. The disagreement can be attributed only to a lack of quantitative knowledge of the aerodynamic parameters rather than to any large departures from the assumptions on which the methods are based. The aerodynamic uara-aeters believed to be princi- pally at fault in the P-IlOK results are dCm/da and and df/da. In the determination of these quantities, data were available from lovif-speed tests made at the Air Tsclinical Service Command, Vv'right Field on a small propellerless model of an early version of the P-i+O series. IJACA ARR I^^o, L5lill In addition, some tests v;ere available rrom the Langley full-scale tunnel of the XP-kO a^id P-4OK airplanes. Data from the Langley tests v/ere somewhat limited since the tests Y:eve conducted for other purposes. The data thai: could he pieced together from these sources indicated not only a large value of de/da but also considerable scatter. In the light of the results given in figure 6, it may be stated that a smaller value of either dCj^/da or de/da vi/ould have resulted in a closer agreement as regards the maxim'om loads at the expense of a poorer agreement in the loads sequence; This reasoning is based on experiences wi-ch computations of this nature (see reference 8) and on the fact that the differential equa- tion of motion on which the calculatioiis are based corre- sponds to tha.t of a forced vibration with viscous damping. For such a system relatively large changes in damping would produce only slight changes in the frequency; whereas changes in tne factors influencing the restoring force - that is, dc/da and dCr-i/da - would change the fi'equency. Closer agreement would result In this par- ticular case if either or both dCVi/da and de/da should be decreased simultaneously v;ith an increase in the damping factor k; and/or the radius of gyration ky The increase that would be required in these factors to obtain a close agreement would have to be larger than could be attributed to possible inaccuracies in these quantities. In figure 7, for the XP-51 airplane, poor agreement was obtained between measured and calculated wing loads • in spite ox" the fact tiiat extensive wind-tunnel data were available for this airplane; whereas in figure 8 closer agreement viras obtained. Figure 1(a) shows that large speed changes occurred with the pull-out shown in figure 7> and the corresponding attitude changes were probably large. Figure 8 indicates that better agreement in acceleration increments was obtained in a relatively short-period pull-cut than in the pull-out represented by figure 7j which required 16 seconds. The experimental accuracy of tail-load m.easurement was relatively poorer than for the SE2C-1 and P-aOK airplanes (figs, 2 to ij. end 6 ) o The agreement shovi'n in figure 9 -or the ET-9B air- plane is only fairly close in spite of the fact that com- plete aerodynamic data vvere available for the actual 10 NACA ARR No. L5H11 airplane and tsll surfaces from teats made in the Langley full-scale tunnel. All the flight tests available from which a pull-up could be chopen, howevei--, v;ere of such a nature that large changes in both speed and attitude occurred during the maneuver and on this account poor agreement might be expected, itlso the flight tests were conducted ac a center-of-gravity location and speed such that the measured tail loads v. ere relatively small. Figure 10 shows the comparison between calculated and measured increments of acceleration and tail load for the B-2J]D airplane. No conclusions can be drawn concerning the lack of agreement in the curves for tail- load increment because of the sparse strain-gage installation used in obt,aining the tail leads. In the interpretation of the flight results to obtain tail loads, it was necessary to estimate both chordwise and spanwise load centers. Errors in the estimation of these centers would cause errors in tail load in individual runs that are even larger than those previously listed. CONCLUSIONS Comparisons have been made of measured and calculated loads on the v/ing and the horizontal tail in pull-up maneuvers for six airplanes ranging in weight from n-jJOO to 1+8,000 pounds. 1. The agreement between calculated and measured wing- and tail-load increments for a specified elevator deflection was good when reliable wind-tunnel data were available and v;hen the airplane maneuvers were in accord with the assTiniptions . 2. Poor agreement was obtained foi" several com-^r parisons. The poor agreement could be traced either to poor quantitative knowledge of the aerodjTiamic paratneters or to the violations of the assumptions on which the ■ method is based. Langley Memorial Aeronautical I,aboratory National Advisory Committee for Aeronautics Langley Field, Va. NACA ARR No. L5H11 11 re?erej:ces 1. Psarson, Henry A.: Derivation of Charts for Determining the Horizontal Tail Load Variation with Any Elevator Motion, N.-Xa aRR, Jan. 19': 5 » 2. Dickinson, H. B.: Ivianeuverability and Control Surface Strength Critt:-^ia for Larje Airplanes. Jour, nero. Sci., vol. 7, no. 11, Sept. 191^0, pp. L69-L77. 3. Perkins. Courtland I'., and Lees, Lester: Maneuver Loads on Horizontal Tail Surfaces of Airolanes. AAP TR No. 14.852, Materiel Div., Ar-my Air'porces, Nov. 19, 19II2. h.. Parkins, Courtland D.: Non-Diruensional Chart Method for Computing the Maneuver Loads on the Horizontal Tail Surfaces of Airplanes. AaP TR No. 1+925, Materiel Connnand, Army Air Forces, May I3 , 19l4-3 • 5. Sllverstein, Ahe, and Katzoff, 3.: Aerodynamic Charac- teristics of Hoi'izontal Tail Surfaces. NACA Rep. No. 688, 191^0. 6. oilverstein, Abe, and Katzoff, S.: Design Charts for Predicting Downvash Angles and Wake Characteristics behind Plain and Flapped Wings. NACA Rep. No. 6I18 , 1939. 7. Street, William G., and Ames, Milton 3., Jr.: Pressure- Distribution Investigation of an N.A.C.A. OOO9 Air- foil v/ith a 50-'fercent-Chord plain Flap and Three Tabs. N^CA TN No. 75I4-, I959. 8. Pearson, H. A., and Garvin, J= B.; An Analytical Study of Vi'ing and Tail Loads Associated with an Elevator Deflection. ^.^C^ ARR, June 191^1. NACA ARR No. L5H11 12 a O SO 00 q O K> (M o IfN o «J*. :f o t~ KN KN » CO 00 r- fTv ^ rt tf\ so r^ ON o C7V CJN vO CO CO (M UN rH UN vo o M 1 _:t OJ o o r-t I-t d O ^ o O o s XI --^ • c CO "^ XI vo" o OV 4^ ^ r\i (7N o o ir» i3 CO so KN CO CO UN o -t o o KN O Ohi (« jO Jx! i»a CO XI «0 u ^ O O o O ^^s o O K\ O KN o o o o 1 o rr\ ITN o rH CO -d- us O o rj i CTS CO o U-N ^ O fH tH r^ o o rH o cu •H Xi O U) 60 XI ^ V4 irv o o r-t o CO J- o Ov s UN o iTV ir\ OS ir\ o r^ J- -d- UN ^ vo rH O vo O > ^ fM o I-t f-t ■-I o o »<\ O « XI •0 "O tiO x> <• 41 vO Q o O O so vO 1 r-i o OD ^ o CTs UN UN o 00 o UTN ON ITN o -J- f\J O r-l i-H O o -i- O CO •H Q , rg rvj O -H ^ oj o rH i-l rH o O »r\ o G 4 Xi X) ■a tiO ^ l» Vl VO o , a C 4 ■l-> « cf § c S tf 4 rH -o « / •4 t* •-I s / ■a . 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Jd ' 60 I 3Z0 --H -7 \ __. XP'5I 300 ^ "^ r i/qht /O ^ 1 1 ■) 1 A <. ) -i c S - Measured -Uicd in colculationd 150 150 260 240 1 J ^ — -- BT'9B ! \ 1 Z 3 4 S b 7 — rtr -- — ^^ ^ --• :z 1 1 h-Z4V NATIONAL ADVISORY COHHITTEE fOt A£«0««UTICS I 2 3 ■<. 3 (, 7 Time , ^ec Ifl) Airspeed time hi:>i-oriei. Fiqur-e /.-Scale fliahi data for oil airplanes under con^iic/eraHon. % ^/^ '^^ • NACA ARR No. L5H11 Fig. lb ■^ /■ \ -2 / / \ 6B2C-/ / / \ Diu-e 6 / '^ " I 1 Z 1 4 5 6 7 r\ 1 1 / 1 \ Di\/e 10 1 1 y V O I Z 3 A- 5 6 7 ■M^ZC -1 ^ ■^ ^ Dive II 1 I 1 / > 1 X I ^ 1 ■4 3 6 7 \ i .Q -S3 c I -10 1 y \ V PBM-3 J I A. J 5 A ^ r t > 7 r ^ I ... ^ / P -40K \j b t A J, 3 4 A (l •> y -^ K? ^ k. — XP-51 F!i ' ' " gni ^ 38 ^ ^Z 44 46 4& 30 5Z j^ 36 Jd 60 -■4 r 1 , 1 f \ XP-51 / r^ r '" /! V -/t? --4 -,1 / / BT-9B / 1 1 > ^ / 7 3 1 ^ J ~ i -> 7 \ 1 1 \ B-Z4D 1 \ \. / NATIONAL ADVISORY COKMITTEt F0« AtKOmOTICS Z 3 ^ S G 7 T'>ne,3ec (t>)Incrernentol-e.let^ofor-deflecfiOn fime hii>-hnie:i. F/Cjure l~ Concluded. Fig. 2 NACA ARR No. L5H11 6 4 I Is o ■■ n / ^ ^ / \ } / \ / \ \ y I 2 ^§ o ^\ / \ / 'A \ M&j6ured // \ I Cole u Jo ted 1 1 1 1 /\ \ \ It 1 \ II \ \ ll 1 ll 1 // II \ if \ Ij \ NATI 3NAL ADVI SORY CO MMITT EE FO II AERC )NAUT CS z s ^ T/'me.^ec s F/'gure ^- Conipar/:ion between measured ond cal aula fed fa//-/ood ond occelerafion incre/"nent3 during o d/i^e pu/houf in an 632C-I airphne. Dii^e 6. NACA ARR No. L5H11 Fig. 3 ^ 4xicP t 'A / A \ i 'i V ^. / \ f '^ ^ o 1 1 1 1 '\ Measured 1 1 1 ; / \ y 1 // \ 1 ft 1 \ 1 \ ^ / / ^ - cot NATIC 1MITT >NAL :e foi ADVIS AERO ;ORY NAUTK :s 2 3^ ^ F/gure3.-Comporhon beiM^een meo:^ured O/ic/ ca/cu- /afed foJ Hood and occe/eraHon /hcremeni6 dur/'m a di\/e pu/l'ouf in ori 6BZ C-/ o/'rp/one. D/Ve /O. Fig. 4 NACA ARR No. L5H11 4~r I sc -/ ^ ^:\ f \\ // \\ / \\ i! \ / \ y / ^ V 1 Measured \ Cnlr 1 ilnf^H \ ^ |1 I ^§ -/ /I \ // \\ ' \\ /'/ \ / y^ / y 1 / / // I // \ 1 ,/ I / Z 3 -4^ Time, ^ec 3 F/gure 4- Compor/:ion beij^/een measured and cole u/a fed iD//-/oad and occe/eraf/'on /ncremenf^ during a d/Ve pufJ-oul in on 532C-/ airplane. Di^e II. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS NACA ARR No. L5H11 Fig. 1 1 1 \ 1 . 1 / // // // / / / ( \ "^ ^^ > N \ o % HHITTEE FOIl AERONAUTICS K Mea::>ured Calculated > a o vS> \ i < \ 1 / < z o >0 / / 5 z / / / o y / ^ y / / y 1 CO "~'~ --^^ --^^ ^ ■ — — - -y — - " ' . . . o •I ^^ W <^ ^ c -ft 1 I 8 ^ Fig. 6 NACA ARR No. L5H11 On c:' e> \j ~+o Q Q) ^^ f> ^ ^ 1 \' 1 \ \ ^ Meo6ured rnlriiln fpd / /-N / / \ / / \ /' / \ 1 / \ y \ ^^7 \ / / \ / / / 1 / ( / / / / >^^ / , '' / / NATIONAL ADVISORY COMMITTEE FOB AERONAUTICS / -r- ^ Figure t)-Compori60n bef^een meoduned ond co/cu- fafed fo/Hoad and aeaeleraf/on increments dunng Q diye pull-out in a P-40K airplane. NACA ARR No. L5H11 Fig. 7 Q) ^ 1 / / / / / / / / / / / / / 4AUTICS 1 / > / / > a j • /■ 3 e AEROI /' / / ' / 2 E FOR 1 / / / / 1 C P 4 MITTE t \ \ \, \ 8 y" ,^^^ ^ C .-- '' - -" 5 ■■■ X \ \ \ \ 1 1 \ \ \ 1 \ \ \ 1 1 \ \ ^ Cak \ fc \ N \ \ 1 >V ■< V "^-^ ^^-- •> i ) ^-'■s ^ ^ Q> ^ ^ p CO ^ ^ CO V ? S '(Ar^U9U/PJ0L// o >< E <5 .^ I B c §. 5 >J o o Fig. 8 NACA ARR No. L5H11 6 ^ St) ^ ^ ■^ 4 .§c^ 2 Qi ^ 'O <{ -KT ^x/O C ^ L Z <;)^ .c : ^^ 1 ^ s^ \ ^ ^t 1 / \ / \ 1 t \ \ / \ \^ ^ / s -/ r \ 1 1 1 -Measured/ \ — Co/cu/afec 1 1 1 i 1 \ / L \ s i \ \^\^ V ^ NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS z T/me ,^ec FiQure 3. - Compor/son defn'ee.n meo^ured onaf co/cu/afec/ f'o/7-/ood ond occe/eroi'/'on /ncre- menf6 dur/no a d/i/e pulhowt /n an XP-3/ a/rp/one. n/ahi- /O. NACA ARR No. L5H11 Fig. 9 c I 3 ?^ ^ / ^ / — ■~-. s / / V \ \, \ \ \ y 1 1 i 6xl(f ^ r^ M , ,\ nauoufGU] 1 ^ \ - r ni^i ,1^4-^^ 1 ■ 1 X ^cjn ^UlU l^U J-^ /"" ^ 1 I "~— 1 1 1 1 NATIONAL ADVISORY k 1 / CO MMITTEE FOfi AERONAUTICS w / / 1 \-^- y / Z 3 4 T/me,6ec 5 r/(^ure 9.- Comporhon bef^een measured and calculated fail- load and acceleration increment during a pull-up in a 3T-9B airplane. Fig. 10 NACA ARR No. L5H11 1.6 I.Z .d § % - -^ 1 \ \ \ N \ / '/ \ \ / \ \ \ \ \ \ / / \ \ \ , ^ \ / \\ / \ \ \ \ \ \ \' ^i4^IO' Z e ^ -z ^ " . — ■ — -^ \ 7^ / / \ \ / / \ \ / V 1 / / ' / / \ ^. \ / / / — ^ ^ MeQ^>uf t^u / / Calculated 1 y Z 5 4^ T/'me, ^ec 5 NATIONAL ADVISORY COMMITTEE FOB AERONAUTICS F^fgure /O.-Comporbcn he.i'/^een meaiured ond caiculai'ed fo//-/oad and acce/erof-/on incn^/nf-n-b during a d//e pu/l-ouf'/n o3-Z4D a/rp/ane. UNIVERSITY OF FLORIDA 3 1262 08106 466 8