1 - UNT Digital Library
Transcription
1 - UNT Digital Library
. —.— ——. —.—— ——-— . . —— — ..— REPORT No. 364 THE PRESSURE DISTRIBUTION OVER THE WINGS AND TAIL SURFACES OF A PW- 9 PURSUIT AIRPLANE IN FLIGHT By RICHARD Langley Memorial V. RHODE Aeronautical Laboratory 6S5 — REPORT THE PRESSURE ~0, 364 DISTRIBUTION OVER THE WiNGS AND TAIL PW-9 PURSUIT AIRPLANE IN FLIGHT By SURFACES OF A — RICEAItD V. RHODE SUMMARY involved in the loads that come into play in order to produce an acceptable airplane, but must only Iinati The investigation n?ported herein toa8 conducted at how to apply the design rules imposed. Langley Field, Vs., by the National kid~isory Committee These rules (References 1, 2, and 3) have prcwen for A6ronaufic8 at the request of the Army Air Corps fo themselves satisfactory, in genend, when appIied to determine (1) the magnitude and distribution of aeroairphmes of conventional type and purpose. As dynamic loads orec the wings and tail eurface8 of a applied to new airplanes of 1sss conventional type, or pursuit-type airphe in the maneurer8 likely to impo8e to new airplanes of conventiomd type but considerably cmlical [oad8on the rariou8 8uba8sembliesof the airplane advanced performance, the rules are sometimes not 8tructure, (i?) to etudy the phenomenon of center of satisfactory in all respects. This is usually not dispre8wre morement and normal force coejlioientrariation covered, however, untiI a structural failure occurs. In in accelerated$ightl and (3] to meaeure the normal many cases it is not discovered at all, fedure having accelerations at the tinter of grarnty, wing-tip, and tail, been avoided by a built-in stxength in escess of that in order to determine the nuture of the inertia force8 acting required. m“nauftanew81ywith the criticul aerodynamic loads. It is perhaps needks to say that crashes resulting The inre8tiga#ion compri8ed eimufianeow meaewefrom structural failures in the air, even though relament8 of pre8sure at 1$0 8tatiomsdistributed orer the right tively rare, have a particuHy bad efleot on the morrde upper %<ng,left lower wing, right horizontal iail 8urfac48, of flying personnel (with some notable exceptions) and and complete wrtical eu.rfacea in one installation and on the attitude of the public toward aviation, and must the came number of poini8 di8iributed owr tho8eportiane be eventually eliminated if cordidence in the sirplane is of the u~ng8 in the 81ip8treamand the left horizontal tail to became deep-rooted. It is manifest, therefore, that &aces in anothm installation, during a ~erie8of lerel design of airplanes must be put on an i the structund j?ight nm8, pull-ups, roU8, 8pin8, dire%, and inrerted indisputably sound basis. This means that design jlight maneurer8. Measured also were the acceleratiana I rules must be based more on known phenomena, mentioned abore,amgularcehxMe8, air speed, and control I whether discovered analytically or experimentally, and positions &ultanew81y with the pre8wre~. less on conjecture. The results obtained throm light on a number of imporWhile a large number of papers have been published, tant queetwne involving 8tructural design. Some of the both mathematical and experimental, dealing with the more interedirg rewlt8 hare beendiwuased in some detail, dernal loads on airplane structures, these hare not but in general the report iefor the purpo8e of making this been correlated to the point where a clesr picture of collection of airplane-toad data obtained in jlight arailphenomena ooourring in the difkent conditions of able to tho8e interested in airpkme 8Wcture8. flight can be obtained, if, indeed, it is possible to do so. The most extensive single experimental investigation INTRODUCTION that has been made is probably the pressure distributestson the MB-3 airphme in 1923 (Reference 4). tion Granting that a major factor contributing to any These have been criticized on the grounds that the increase in airphme performance is a deorease in weight, airplane vras of a ~ery speoial type, and had individit is cIear that since the structural weight of an airplane ualities of such nature that the results were not applicaconstitutes 20 per cent or more of the total, any saving ble to the general prcbkc.. While some of this that can be effected in the structural parts is worth criticism is well founded, it is useks to expect, except while. But to design a structure light yet safe, the to a limited degree, that complete pressure distribuengineer must have a through and accurate knowledge tion inves&~atione on any airplane -will furnish data of the oharacter of the loads that his structure must suitable for the solution of any particular prcblem. withstand. Actually, of course, the designer need not This is true because any airplane is neoessariIy individbe thoroughly conversant himself with all of the facto= 6S7 — .= -- .— — — .— -—— -= .= ..-. - .— .-— — .. 688 REPORT NATIONALADVISORl COtiMTT?iT)IU FOR AERONAUTICS ual (unless duplicated to a very fine point of perfection) and also because, which is probably more important, the labor involved in these investigations is so great that it is impossible to treat Rriy one phase of the Ioad problem adequately if a fairly complete picture of the whole is to be obtained. , distribution. The results are of immediate interes~ to those agenck responsible for structural design -, ~d even though not analyzed to any great extent shouId also be of value and interest to airplane designem. To expedite its presentation there has been no attempt to analyze completely any one phmo — FIGUMI.-Front viewofITf% alrplana Thus, the present report attempta to portray the phenomena occurrigg on a p.umuit-type airplane in the maneuvers that it is called upon to perfow, or what amounts to the same thing, in special test maneuvers outlined to impose the same conditions of load that recur at the critical times in the more familiar of the structural phenomena that am brought to light. Instead, the present report presents the data as obtained, worked up to the stage where they can reridily be used in studies of design methods, for the consideration of those concerned, and has in a few instances called to attention the more obvious points in which FIGURE Z—TbIe8quart.ar front vfew of PW4Y airpkne maneuvem. To this end, pressure measurements were made on the right upper wing axtended to include portions affected by slip stream, fuselage, and windshield, the Ieft lower wing, and the tail surfaces of a with amelerBoeing PW-9 airphne, eimultaneoudy ometer readings at the center of gravity, wing tip and tail in the maneuwm above mentioned. The data obtained represent a ve~ extensive collection of information on structural loads and their structural design methods now in use me open to criticism in light of these redts. The flight tests were made at- the LangIey Memorial Aeronautical Laboratory in 1927 and 1928, at the request of the Army Air Corps... APPARATUS The airplane,-The a slightly moditied airplane used in these tests was Boeing P177+ pursuit airplane. .. . . . PRESSURE DISTRIBUTION OF A PW-9 PURSUIT AIRPLANE LY FLIGHT 689 (I!lga. 1,2, and 3, Table I.) The miIitary load, irdudI and also noticeable @ Figure 10. The effect of this ing the main tank, -ivas removed and the top coding gap will be mentioned in the %ussion of results. Pressure orfllces and tubing.-InstaUation photoforward of the cockpit raised sIightly w that the test graphs are gimn in F~ures 7 to 11, inchsive. The instruments and apparatus could be accan.rnodated. The pressure tubes leading from the upper vzi~u to ori6ce and tubing installation is essentially the same the fusekige formed faIse struts which increased the as those used on previous tests, with aluminum tubes used throughout., except for short and easiIy replaced drag and lowered the high speed about 6 m. p. h. On lengths of rubber tube at the manometer connections the other hand, the weight was reduced 50 pounds and between the fixed and movable surfaces. A and the stalling speed lowered about 1 m. p. h. The diagram of the orifice locations is given in F~ge 12, pilot reported poor directional control, but the longiand the type of ofice used is ilI@rated in Figure 13. tudinal control and aileron action vm.re e..cellent. On Manometers.-The oritlces were connected to two the whole, therefore, the performance and maneuverN. A. C. A. type 60 recording muMple manometers ability were not reduced suficientIy to affect the which were located just above the center of gratit y significance of the rew.dts. in the space formerly occupied by the main gasoline The wings of the F’W-9 employ the GWingen 436 tank. These manometers are the same in principle airfoil section throughout the span. They are, howas those used in the MB-3 and ~~–7 tests (References ever, tapered in plan form, and the upper and lower wings dfier in plan form from each other (fig. 4); I 4 and 5), differing mainly from them in that they ..— .— .- — . .— FIGURE3.—Three-qmrter rear view of PIT-4 sfrpkm in addition, the upper W@ is washed in at the center Table II gives the actual ordinates of aU section. sections at which pressures were measured. During the preliminary tests the airplane broke a wheel while taxying in after landing, pitched over on its back and damaged the central portion of the leading edge of the upper m@, which remained ahghtiy deformed subsequent to repair. OutIines of some of the deformed ribs are shown in Figure 5, compared with the true sections. A structural feature of this airphme which has a bearing on the results obtained in the tests is the lack of flying wirw in the rear truss, redting in a structure 1sss r@d in torsion than the normal single-bay biphne structure. This dews the incidence of the ceLhde to vary with changimg load, pmticdarly in low angle of attack conditions of flight with the center of pressure well back. Another characteristic of the sirplane that shouId be mentioned is the slot-shaped gap between the wing and aileron, illustrated diagrammatically in Figure 6 accommodate 60 pressure units each instead of only 30. A diagrammatic sketch of the attachment of a-pair of orifices to a pressure unit or ‘tcapsule” is given in F~e 13. Other instruments.-In addition to the manometers, the fobxing instrnmen ts were used: N. A. C.?.A. recording air-speed meter ~eference 6); N. A. C. A. recording turnmeter (Reference 7); NT.A. C. A. control position recorder (Reference 8); thee single component accelerometers (Reference 9), located as shown in Figure 14; and a timer. A movingpicture camera was also used to measure an@e of attack as will be exphiined later. 31ETHOD The method used in thwe tests does not ditler in any essentird feature from the methods employed in previous pr=sure-distribution tests. As it was desired to obtain results for all of the ying and tad surfaces simuhaneoudy, if possible,’ the orifices were disposed to cover the right upper wing, left Iower wing, vertical .— — — .—. 690 REPORT NATIONALADVISORYCOMMITTEEFOR AERONAUTICS tail surfaces, and right horizontal tail surfaces. Limitations of capaoity of the apparatus that could be carried aboard the airplane prevented the investiga—. right upper and lower or left upper and lower) was imposed by the impracticability of running all of the pressure tubes into the fuselage from one side, and also i6 ‘O” [ —4z”— .-. 150” + 1 ..O ~ 4 .1 ‘o Q I 1 aivodl - . —- 1 ~ titer , strut - t merits 1, .1 I 1 I I 8 1 i t 1 t.. I U-J. a ? 9 . ‘..1 A ...-= ,— .— Lower whg A Yxw -— v c SfobifLzer & de vato~ a, Stay wire a f kmbmefits FIGURE 4,—PW+3 wings and tall mrfaceswkh sparlmstlom tion of the remaining surfaces simultaneously with those mentioned, but an independent set of rune was made later to investigate more thoroughly the slip stream section of both upper and lower wings and the left horizontal tail surfaces. The right upper and left lower wing arrangement (in place of the more desirable — ‘Hinge & Fin 42rudhr and strut attachments because it was not desired to unbalance the airplane even slightly unless absolutely necessary. The pressure measured at each point was the algebraic sum of the pressures on the upper and lower surfacas (see fig. 13), no attempt being made to measure the pressures on these surfaces separately, except in a PRESSURE DISTRIBUTION OF A PTV-!3 PURSUIT AIRPLANE 691 IN FLIGHT few cases after the main tests had been completed. I ground effect. It was feh that level horizontal flight Pressure curves were mechanically itttegratail to obtain 1 could be maintained accurately enough in this way to allow of the use of inclinometer readings directly +s loads and centers of pressure. Simu.ItaneousIy with The method failed because the slight the presmre measurements, records were obtained of angles of attack. variations in engine speed and wind veIoeity caused air speed, normaI acceleration at the center of gratity, left upper R~U tip and tail, angular velocity in pitch I the inchnometer to oscillate enough to make the readinge quite erratic. Further attempts were made, flying or roLI, depending on the maneuver being investigated, as before, with a moving-picture camera mounted in tind control position. These were synchronized by Dw. ..—— .— — .— 0 . ----- __ GZHingen -. —-—------ 436 _-.- — .— — -—— .— -——— ._-— FIGUEES.-Comparison of rib A, B, and C with trus OdttIogen 436Se&on means of the timer, and all of the measurements made were plotted together against time to furnish a history of each maneuver. It will be noted, in glancing at the center of pressure data, that resultant centers of pressure are plotted in terms of per cent of “centric chord.” The “centric chord” is here defined as the chord passing through the centroid of the plan form of that portion of the wing extending from the root to the tip, and it is used instead of some other arbitrary datum, because the position of the mean C. P. on the “centric chord” of tapered wings corresponds fairly closely with the mean C. ~. on the constant chord of straight wings. With respect to the maneuvers in-mst~~ated, speciaI attention was given level flight and pull-ups. A considerable number of leveI flight runs were made in order to furnkh a basis for the study of the results obtained in accekrated fl@t, and also to furnish data for comparison with wind-turmeI results. Attempts to measure angle of attack in leveI flight faihxl. First attempts were made with a pendulum inclinometer mounted in the cockpit, the pilot flying horizontality close to the ground, but with a sufficient aItitude to eliminate possibilities of encountering 41630-31~ / ‘ , I the cockpit, the lens axis being normal to the X2 plane. vertical reference lines on a row of hangam were photographed, and the wgles of thwe lines on the picture with the frame ~Ue were taken as angles of attack. This method showed promise, but with the hangars located on the south side of the field, as they are, it was impossible to obtain c.lear pictures, inaarnuch as the verticaI reference Iines mentioned above were aIways in the shade. hgles as obtained were probably correct b within 1°, but this accuracy was not sufficient for the purpose for which they were desired, and all of the angle readings were thrown out. Centers of pressure, therefore, were plotted against normal force coefficient as the independent variable. The extensive investigation of the pull-up was made for severaI reasons. First, it is one of the few maneuvers that can be subjected to reasonably close contrdj or repetition with accuracy. In other words, it is a simple maneuver requiring the piIot’s attention on only the initial air speed and movement of one control surface. It is, therefore, possible to obtain a graduated series of maneuvers to allow of the study of accelerations, anggar velocities, etc., as afhxting the distribution of load. Also, the pull-up shows the &- — - — -— —— —- -- 692 REPORT NATIONAL ADVISORY COM?sHTT.JIEFOR k“ AERONAUIWS “ I FIQUM 7.--Side vfewofPW’4 afrpkmewith hstrnmmd panelsremowd , FSQURE8.-Detnil view of ms.fninstrument instdh$ion showfng tuk+ connectionsto manometers . PRESSURE DRTFKBLTTION- OF A PW–9 PURSUIT AIRPLA3E IN FLIGHT 693 FIGUZS 9.—DetW vfew of acceIaometu LustnnatfoItInt8n Show@ alK)mkmti~ ● I FIG~X 10.—Wing tip aeeekrometeirLnshIMon 694 REPORT NATIONAL ADVISORY tribution of pressure through a large range of angle of attack and furnishes direct information on the most important loading condition, viz, high angle of attack. Third, the unusualIy far forward position of the center of pressure at high angles of att8ck in accelerated flight indicated by tests on the WZ-7 and Tiil “i@lanes (Reference 5), as well as the coincident high values of normal force coefficient, made it desirable to study the high angle of attack condition at some length with the hope of discovering reIations which might account for the phenomena noted. Because of the importance of accurate air-speed measurements in ~obtaining normal force co~ieqts, the air speeds as recorded: from a Pitot-static head mou!nted on the front outer strut were carefully calibrated against those obtained from timed runs ‘---’-——- ‘-- ‘ ‘----“ ‘—-- eluded there that individual pressures arc correct to tithin + !2 per cent, while values of load aro correct to within +4 per cent. “Recent investigations of the effect .of temperature on the capsule calibrations indicak, however, that those figures should probably be incfeased by about 50 per cent for the greater part of these tmta, and in some few cases should be doubled, These errors, however, have a minor ~ect upon the measured distribution of load, since the temperature erroti ire, in the main, a certain percentage for a given temperature regardless of pressure. Furthermore, most of the capsules, being identical in construction, show comparable temperature errors, Air, speeds are correct to within 3 per cent at all spee~ in level flight. They are probably correct withig. about 4 per cent in accelerated flight, sinco it FmvE~ 11.-Detd.l oforifk?afnstallntlonin wing over a measured course, and also those obtained from a suspended “bomb” air-speed he~d. (Reference 7.) This calibration, it was found, sufficed only for level flight runs and the initial air speeds of the maneuvers. It was necessary, for the pull-ups, to metm-rre the air speed by another method. For this purpose an airspeed head was mounted on an outrigger about 5 feet out and slightly forward of the lower W@ tip in order to eliminate interference from the wings at all angles” of attack encountered. ~ speeda obtained from this head checked the~(bornb” readings within one-fourth of 1 per cent at all angIes in level flight. It was thus considered that readings obtained from this head in vertical plane maneuvers woild be satisfactory, and all of the air speeds for the abrupt, power-on pull-ups were, therefore, corrected on the ‘basis of results obtained from the outrigger. head. PRECISION A discussion of the sources of error in. pressure measurements, using the methods applied in the Tresent tests is given in Reference 11. It was ccm- is believed that interference effects have been largely eliminated, although some uncertainty still exists on this point, Individual values of rib center of pressure are correct to within about 2 per cent in the high angle of attack conditions, except in the case of rib L on the low~wing, and within an increasing error as the angle of ‘attack decreases until, for conditions approaching zero lift, they are quite erratic rmd unreliable. For this reason, momente instead of centms of pressure are given for these last cases. Imgituciinal centi-m of preisure of” resultant forces are correct to within + 1 per cent-at high angles “of attack, while lateral centers of pressure are correct to within about 2 pm cent. Accelerations are correct to withh + O. 2g except where noted. Control position angks are probably not correct to within less than 2 or 3 degrees, since the instrument was connected to the control levers in the cock: pit, and did not measure change of angle caused by deflection of the control system under load. PRESSURE DISTRIBUTION OF A .PW-9 PURSUIT AIRPIu4NE IN FLIGHT 695 — —— .4 . — - .—— -- — *per, wing < —— -.—— =— - - . . —.— 35 “ Stabilizer’& Fin g rudder elevaf= FIGUEX lZ.-Lccatfon of orificeson wing and Ml surfacei Location of siatiom .=---- — .— .— - ,= ~ .— 696 REPORT NATIONAL ADVISORY COMMITi’EE FOR AERONAUTICS FIGURE13.–Dlagmm showkig type of onflw and connectionto capsule I “C, C. G. occekrotnefer W, Win occejerome+er T, TOI“? occeleromefer 7, . . . . Pw-B oirplone showing uccelerome+er locdions ._— t FIGURE 14.—Thre8-vlswdmwfng of PW-9 airplane sbowlng nccclemmeterkxatfom PRESSURE DISTRIBU’IXOX OF A PW–9 Time synchronization is correct in most cases to within about one-twentieth of a second, although in some rum, because of instrument diflicuhies, the synchronization is rather poor. For this reason, any calculated quantities depending on the records of two or more different instruments maybe quite umeliable. This is particularly true of abrupt maneuvers in which the measured quantities vary tkwough a wide range in less than 1 second. PRESENT~TION AND DISCUSS1ON OF RESULTS The results following are presented and diswussed in -groups according to the maneuver under in~e@ation, tiz: 1. Level flight. 2. Pull-ups. 3. Rolls. 4. spins. 5. Irwerterl flight. 0. Dives. i. Pull out of dive. In addition, the tail load, slip stream, and fabric preesure data are summarized and presented separately. Yi’bile alI of the data obtained in the tests are not gken in their entirety, representative examples of the most important cases are included, and sIsc, where it was felt they viould be of interest, more complete data are given. Level fight .—The leve~ flight results are gken in Table III and in Figures 15a to 20. Figures 15a, 15b, 15c, and 15d show the distribution of pressure in terms of q for four representati~-e cases through the speed range or usefuI range of engle of attack. An inspection of these figures at once discloses sereral salient points, viz: 1. At low speed, or high angle of attack, the center of pressure locus is, for all practical purposes, at. the same per cent of chord aIong the span untiI the tip is approached, where it bends suddenly to the rear. This is particuhdy true of the upper wing. With respect to the lower ving, this point is queationabie since pressures at only four points were measured on rib L, and the accuracy of the pressure curve at that station is poor. 2. As the speed increases, the center of pressure moves back varying amounts at different stations along the span, the trend being farther to the rear as the tip is approached untiI, at high speed, the center of pressure at the tip is almost twice as far back as it is on the inner portion of the wing. 3. In practically all cases, minor peaks of pressure occur od the aileron. This rearward trend of the center of pressure from the center Iine ta the tip at low angles of attack is Figures 17a to 18e afford a worthy of special note. closer study of the center of pressure movement. Figures 17a and 18a show the nriation of resultant center of pressure with norma~ force coefficient for upper and lower wings, respectively. Figures lib to 17h and PURSUIT AIRPLANE IN FLIGET 697 18b to 18e give the center of pressure vs. CMfor each individual rib. As might be expected, the points for ribs in the slip .strema and near the w~u tips are somewhat erratic, but a suilicient number of runs were made so that fairly good curves can be drawn in every case, with the exception of rib L. The rearward trend of the center of pressure toward the tip as normal force coefficient decreases is clearly apparent from these curves, and leads to the suspicion that a considerable amount of tm-ist (washout) exists, particuhxrly in the upper wing. This suspicion is strengthened by subsequent data which show a rapid decrease in load along the span toward the tip at low angles of attack. It is desirable, therefore, to study the eff~t of twist on the load distribution at some lengt.h. The subject is discussed wry weIl in Reference 12, in which the author points out the importsmce of alight angles of twist at low angles of attack, and calls attention to the fact that the twist may be a budt-in feature of the wing, geometric or aerodynamic, or lit may be a torsional deflection under load. He also gi-ies methods based on the vortex theory for calculating the load curve for wings having certain simple conditions of taper and twist, comparing the results with load curves computed on the basis of the “strip” method, and pointing out that the laiter method is not sufhiently exact in some cases. The method is given more completely in Reference 13, and is extended b include any tapered wing with any character of twist. For present purposes, it is sufficient to compare the load curve obtained in fLight with load curves calculated on the basis of the strip and vortex theories. TO do this, it is pecessary to choose a particular condition of flight and to determine what the actual twist *SS in this condition. The condition chosen is highspeed le~el flight, since sewd runs are a~ailable for & condition, and hence experimental errors can be To deterreduced by taking an average load curre. mine the twist of the wing in this condition is not difficult. The most reasonable way would seem to be by actual measurement of the rigged incidence and adding to this the extra twist induced by the application of the load existing in the particular condition beiig studied. Unfortunately this method is not possible, since there is evidence that the rigged incidence changed during the course of the tests., and hence the true rigged incidence for any particular run or set of runs is not known. It is necessmy, therefore, to resort to a di.flerent method. Assumptions are made as follows: 1. That the curves of C.. -is. a for individmd rib sections mot too near the tip and outside of the propeIIer influence superpose upon one another; that is, that for ribs B, C, and D for instance, the center of pressure is at the same per cent of chord at the same 2. That scale effect does not inangle of attack. fluence the center of pressure position except to a negligible extent. With these assumptions, probable ---— .— ..— -—— — .- — ..— .= — — _ -- — . . -— — .— -– — L FmuBm16s . , BIounba s-mum 16d 1 1.0 1.0 ,8 .8 o m ,6 .6 Cw 10 a 6 Fee! along 3ern?spon 2 FIOUEXIa-varht!on ,4 .4 ,2 .2 00 2 4 6 e Feef along of normal r0r09coemotentalong the mu in level ftkht 10 sernispon f2 /4 16 throu@ theWMMIrmm 02’ co co . 1) 20 -— - L@per wing, Cl? referr to ~nfric &or-d 0[ ed Rib A . I . . Rib B Rib c . . 0 a> 0 =ir ~ ~ ~ Ed H ~ — 4 0 6(.I , I(a) I ,, 1 ml 1 I I(b) I II II II II(c)IIIII III l(d)lll]~ z * E $ &30 ? a ,. . . . 3 o ? ~o ~ tb ‘ —- Q :20 Rib D : ti~ Rib E a / 4 0 40 0 ● 0 A ~ ~ , Rib F — r 9 m 0 , . / (e) (f) (a [ 80 ,4 .6 .8 0 f 4 .6 -8 1.0 0 .2 4 c. Floo’rm17–Can&r of rumure in w E D i ‘ I .2 ! 1~ 1 .2 ~ 0 (h) 1! / /oo0 0 > 0 ; 0 0 — — — — — — — — In 0 E Rib G - 0 60 * c1 cant of chcrd m, norml form roullchnt -.6 4 “.8 o 2 .4 .6 .8 PRESSURE DISTRIBUTION OF A P’iT-9 PURSUIT AIRPLANE IN FLIGHT 701 .— — . -. ~ ..-— — .— ..= .— . . - .- -- .= .— .. — —— — 702 REPORT ] I I I I I NATIONAL II II ADVISORY COMWFTEE II I I I $0 ——————–- \ !? m-~ I ii w 1, cl [ !11 t D I FOR AE)30NAUTICS I I I I ‘. Tokwon”a/ de flecfim I I I I I I I I I I I I 11 t’~-.? % ~- -3 $ 0 2 4 /0 6 ..8 Feef olong sernkpon !2 FIGUREN.—Torslona1 deRectIonand total twfst In h&h-s@ Feefolong FIGURE21.-Compiu’f.wn of experimental loedcum /4 . Iavel fflght semispan and loadcarvesbawd on stdp and vortextheories /6 . PRESSURE DISTRIBUTION OF A PVf’-9 PURSUIT values of center of pressure for ribs B, C, find D in high-speed level flight are determined from Figures 17c, 17d, and 17e. These values are referred to Fwure 19, which shows the variation of C. P. with a on rib C of the PW–9 model, and the values of W@ of attack obtained therefrom. The differences in these angles are taken as the twist of the wing between the stations used. Referring to Figure 20, these diHerences are plotted against span, taking B as a starting point. The curre is then disphmed downward until !ts inner end, extended, intersects the base line at 3?4 feet, the point at which the wing chord is a maximum and which is subsequently to be taken as the root section for the comparison of load curves. The outer portion of the curve is extrapolated to the tip as shown. This extrapolation is not particuhdy hazardous, since it is based on a curve of measured rigged twist (not shown) which has a similar form, diftering only in degree. Figure 20 represents then the twist of the wing in high-speed level fright, and includes both the ~~ged twist and torsional deflection. Using this twist and an angle of attack for the root section based on the a~erage rake of CN at rib B and the slope of the CN vs. cc curve for a similarly located rib on the model, the span load curve is calculated on the basis of the voruxx theory with the resuh shown in Figure 21. The agreement is quite good. The desigg load curve shown for comparison is based on the strip theory without Wowing for twist. It is only fair to point out here that the rigged twist cm this particular wing is rather unusual, and is greater than the washout required in the design and specified by the manufacturer to compensate for propeller , torque. The discrepancy between the design load curve and esperirnental curve is-l therefore, greater than it should have been. But, further in this connection, it is neoessary to mention that whiIe the torsional deflection under the applied load in this case ody amounted to one-half of a degree at the tip the appIied load factor was 1, whereas the condition designed for at low angles of attack is the early stage of the pull out of a dive, in which the apphd load factor may be several times unity and the torsional deflection, therefore, increased in proportion. It is manifest then that not onIy is the rigged incidence a matter of importance to ccmider in the. structural dwign, but the torsional rigidity of the structure as weII, and this latter quite apart from considerations such as wing flutter and eeconda~ stresses. This conclusion is not new, but seems to have been overlooked h this country. In spite of the probability that stress analysts and designers wilI tie-w with distaste the necessity of carrying out the additional calculations required to obtain a more exact load curve, there seems to be no simpler alternative at present. It may be remarked, however, that the solution of the load curve, once the twist is known, is neither dif3icult nor AIRPLANE IN FLIGHT 703 particukdy laborious. In relatively flexible structures, however, because of the mutual @ationship of twist and load, it becomes necessary to resort to trial and error, always a tedious process. The span moment curve is readily determined on the basis of the assumption that the moment coefficient varies Iinearly with the Iift ccefikient when the moment curve for the basic airfoiI section is knovmt. The minor peaks of load occum%g on the aileron are believed to be due largeIy to the efhwt of the slotshaped gap between the wing and the aileron, since the control position record showed that the aikmn was neutral or very near neutraI in aII runs. Pull-ups,-Pull-ups were made as folIows: 1. Abrupt, power on, through the speed range. 2. Graduated abruptness, power on, at 70, 100, and 130 m. p: h. 3. Repetition of some of the foregoing ~th power off . Figures 22a to 25f show the variation in pressure distribution throughout four abrupt pull-ups at speeds between 79 and 181 m. p. h., and Figures 26 to 29 the corresponding span load curves. In the puII-up at 79 m. p. h. the airplane was nearly stalled before the maneuver was begun, and the pr~e curves for O time are, therefore, characteristic of the high angle of attack condition in steady flight. As the maneuver progresses, however, the angle of attack of maximum Iiffi is reached and the mean ceni%r of pressure (see @. 30) on the upper wing movss forward to about 26 per cent of the centric chord. It W be noted from the pressure plots for the medium and high-speed pullups and from the time history curves (figs, 30 to 55) that the maximum forward position of the center of pressure is the same within the experimental error in every case, the average value being about 27% per cent. (Exceptions: Rti 73 and Run 137.) This vahe agrees with the vaIue found in wind-tunnel teds on the PIT-9 celhde. @.ference 14.) It does not, however, check the value found in monoplane tests, the ditTerence being about 3?4 per cent with the fuUsoale center of p~ure farther forward. This discrepancy is ascribed to “biplane effeot” or the mutual interference of the upper and lower wings. On the lower wing the center of pressure do- not move so far forward, reaching an average value of about 31 per oent. This value does not check the wind tunnel value, but comparisons are not strictly valid since the wind tunnel Iower wing model was extended to the phme of symmetry, and hence has a greater span, proportionally, than the fuU-scaIe wing. While the iduence of biplane effect on the center of pressure is not an unknown phenomenon, thus far it has not been taken into aoommt in the design rules. The itnport.anco of this phenomenon should not be overlooked. While it can be seen from F~res 56 to 59 that the distribution of load aIong the span on both . ..— ~ — .— .- .— — — — .—— .— .- .~ ..— — — .— . .--+ .— ..— . — . .— mom FIOUES22u rmum 2a .—-. . . . Z211 rmmum . ,,, 4 ..— — Run 65: ot 79 m.p. h. .. -., -. .. 1– —-— . u Time: Y .0 sec. ?ufl FIOVRB m molmn 2X R1OVRB 2% a Flouun I * . ——._ — ..-— — FIOUM28a ,.. Flaul?ll ti ., ! “1 .,, ,. 1 1 Fmvm * ,!” I I —,. . ... .. ..-.—— . .---. —— —- “--” .“-~/$$J .. ,.- “-”””” ., .. IA N \“:”“>,; F ‘“ ;’s,:” \\ -J “ P. 32 ““,+ ~1. ““:i: ;;, “ ,,., .“ -3>...>, . 27 .,%, ‘p g >< Y /Ji”\. /-31 ‘L 27 ~,’\,,,, , ‘I ~oo.g j5~ 1,,) L it 1oo-g$ 50 $4 0“ ~ , \ J ‘. A- d I___ ,,;3,, “ Run 130: ful/-u 116 m.p.h. Time:.f3 sec. .—— .-. —.. . . .-—_—— _____ / \ ‘~:!jti~ % L“. ‘, 1. 0’ ~lv . . ..—.—> .,.— ..- JJIGuEnW — .— FIaunE !2% .—— 1 . Waunn 28g o -! A . I FD30’M 91b ,, ,.:. — -. ——— ——-— lWmRB fl— Zid maum 24a n —.. ..,,. ————— ._ MICIUN !ug * :l”’’il~ ,. “ “ , .I:, j,, j ,, i,, :’jl’ /j,, l,” ; ,,( ,,,, ]i ., . -——— ....—— FmuRm m FIGURE % II ,+,: ;-.4.,,!, ), ’b. ., ..... . . . .. . ,., “ ; n .-. — . ..— — .-.—.-.- ..— ----- ~ . .... —.-..— —_ -------- . —., ------------ .-. — ~ -- --.-–- _,,. , I.---.., ‘ Y’ /“” -1 1- Fee+ along semispon FIGURE20.-V8rlatlon ofspan lead dlstrlbntlondnrlng a pull-upat I I 10 1 .- —-—- 1 I 1 1 i I I 8 6 Fed 01w9 4 semispon 2 FWri /y roof, _ I b opf 1 I I ’’4’’’ I I 79mika per hour. (Run No. M) 1 I 1 , 1 I 1 1 1 1 1 I I 1 1 1 1 1 1 E I 1 Im. L”’’’ L’’’ Feef J’’’ /L’”/! ulong semispon FXQUM‘Z7.-VuWionof span lomldkirlbution durfnga pull-up at 110mlfmpm hour. (Run No. 180) ,! ,,, , 1’ ,: ,,, , “’/9’5 800 800 700 700 600 q 600 ~ SO05500 -. 4oo.f400 1 3004.900 6 2008200 IL70 /00 10 8 2 fCd 0°0 2 4 %Ot7g SCm$7pQ~ FmuRu 2&l.-VariailonOfa~n ]onddiqtrlblltlonduring a pul].up at lfl?mile FIOURE20.-VUWOU G F=+ 12 ulon~semispo~ per hour, (Run No, 182) ofEIMIIload dirntrlbulkmIn n pull-up at IN mll~ par hour, /4 16 o % P (Run h’o. 137) I -1 I I I A Ill I z. I 11- 200 0 -200 Uppe ; 2000 I I wing load -\ I I 1“1 H ?7 l-T- ! --Fill ,&. 1 1 1 1 1 , 1 I I I , . , 1 I.. oil Q.__J_l-L!-+- gl I I ‘—-’4.-L’ 1, I “’;’” ~no *“ load I ‘ 0 1.4 c%.? \ 1.2 1.6 Upper 1.0 wins , I n, S.L cnL I I I I I ‘L. awer ,8 - ~ / / $ I 4 1.2 — — — ~ bz .6 ~ I I I wing I I ,-,- w FmIIRE 811 —Time Mtory of an abrupt G 7Zme,seconds powwa pdkp at n dlm w . hour. (Run No;05) FIOUEXSI.-Tlme fit-netseconds hLWoryd an abrupt POWWOUP*P at W roll= per hour. (Run No. 66) I K 110 \ 1 I Air speed. .,’ \ — 1 v ti QI.I I-. K b,- m / I I I I I 1 l,- I I , L+ 1 41 . I ! ! ! I I 2000 Toil momenr’-l -x 3oTttTlti-tl —.: -1~-i 4 I I I i i MT 1 Upper 1~ wmg Iood, / yom — \ ..”, ,— , .- F 1 g’’000 2.0 I I I I I I I – , , I 7 w /.---f~ii I%u Cf’l&2 1 ●s I u tipper’” wlAg ‘ f.o I-i-+4 1 I -+ !/’! w y 1 J’ I I . I o -1.4 I ., o + Ill – 1 1 o m – – - – – – – =’ /mold . ~ = – u-11 /.0 “-1 / I I I I I I I Uver w/n9+... I I I Lbner Time,seconds FIUIXOIS2,-Time hlatory of an abrupt power-m pull-ap at 109rnlka par hour. (Rim No, 07) I — / M--12 I /’ J $ – .- ~$ij .< (.2 d%.l__Kll / CH / <“ - f. 2 @ L wh- Time,seconds lWunM S8,-Time I. 6 4Mory or an sbrupt porve~n pull-up at 110mllm por hour, u (Run No. M) I 130 — — ti .%i I I 3a *“ EIevator angle’go yQ — L I 1A i, -. I Iv I I I *’2EEEk.8 w . Ttme, seconds Erauu S4,-TlnM hlst~ of an ahnpt IMw-n No, lM) WIMP at 110* x hour. (Run Time. seconcfs UBS&i-.TIma hbtorg or an abrupt ~wtu+m No. 1S1) pull-up ●tlamlh per Imnr. (RUU I ,~ c * 100 I I 1 I I I I /Y I fie,- /.oH—H—H4——H+H— \ POoo / .-l I I Cll w 4$ ]+ 3000 ~ I , I I Kfi I I I I 1 I 1 1 1 1 I I I I I \ I I -1 I IM lTa(lrny+n) ! I Upper $U -20m ,.lm .~ “ / I \l 1 / I I I I / 1 ‘--- wing food - -... I I I I I /.41 / tipper “wlii9 Iood-td1 1 J 1 i“’ ‘ ‘- I 1 I ‘- 1 1 1 1 1 1 r I 1 nI Lower who load.A \ \ F’ltf”2 1, Ci 3000 ~ “x ~20004 1.0 > 6? 1,< A---lo .8 f. t-l /,0 I I I I—I—U—U—L44——U ,, ,8 nk I ‘1 I I L 301 a. o .,ln I I I I -. -.5 L Tth7e,seconds mrc, seconds FIGURE 80,-Tlmo F1auB~ fJ7,-Tkaehlnkry or an alm’uutIMwnr-on pull-up at 140mllm Por hour. (Run No, 188) hlalary or m abrupt [email protected] pull-up at 137mllu per hour. (Run No, 13’2) ,1, ● I I 1 I ,1 ,,1, i ~1’ !’iii;’ ‘ ; 1 ii 11’ I:l,jl j: ,,, “ , ‘IJ’l,[l I i ,’~il’:~lll~~iil.ll,l~l,,l,l I 11111,,; ,, “ . i70 . .X z , , ! I Ele vofor ~- ! fh41 ! cbgte - -. , 1, x I I I I I f30 I h t A I A - 1 I I I I I I A <W A.-m +).- wefe,-;+, -1-vek -aOOwl+H+H+wH 1 1 1 1 I I I II I ~ I I I I I I II -200a ,1 VI I YI A “%’ 1.4 .anw* 31 .a ,@ Iif”/.f@ -, L9 6 I — T301 -2 ! Lww=l -m d / ‘- -- I k Wtl I IH! - Upper I wing L I .8 I I I I wing loo~ I I I v 1. & ,. I Lower I I Upper I I I I I A /1 ., win.q /A .6- ?’ (-Jk /1 I J l/l- I 1/ / / / < 1 1 1 I k, 1.5 \ ‘Lo wer 1 I le.-,- I wirm w;”,-. I Lod H-wl ~ 601 FIauM ,, I I 0 I I I I .5 upper wmg 1 /.0 ‘ I!fos iZme, seconds 8D.-Tima hlsmry of an abrapt pawwon wIf-nP at U.lamRea par Maw. (Finn No. MS) ‘ I-t-tw-ht’-m ‘.i!w-thd+l-,+, I I ‘! er ,I’Low LL7L-’! I py:~’~9J’q\ y Vine, seconds Fmuua40,-Tlme hlsLoryo[an abmpt powar+npuJ1-up at 172mflaspr110ur. (RunNo,184) 1, I ,; ,,1 .,,, ! 1’ ! \l Illi’,iil ,!: ,1] ’1, ‘! 601 1 1 1 0 I I 1 1 I I I wing I “.- w 1 1.0 ,5 Time,seconds Flauns 41.-Tlnre hlutow of an sbmmt ~wara how. (Run No, la7) 11.I-)F9 . 1,5 Pull-up at 181mlho w , 720 REPORT NATIONAL ADVISORY COMMITTIZE FOR AERONAUTICS . . .- L- L T I 1 1 1 I 1 1 wfngt -, LPI- I T I —. -. , ?h7e, seconds FIGURE4Z.-Tfme Irfstory of a mfld powewjn puU-upat 110MUM per boar. (Run No. 78) PBESSURE DISTRIBUTION OF A PW- 9 PURSUIT AIRPU~E I K 60 1 I 1 I I 1 1 I I I 1 I I I Etevafor anqle. 721 LW FLJGH1’ -.= .— -— .- 30 4 i / “ 1 I t I t I I t 1 I 2of I — . ——— — I I — — — Td ucc&e I go I ! I I I 1 1 I , I 1 1 . < $! veloctiy ~-- > .4rgtIti 50 I I Ilu+--r I I I I I I I 1 I - I ~ p I I I 1 ; I I — I -m -lCW . 1. — — -m 3000 -. — I I I I ! t 1 ive w-m bcuitbgs t 1 I /.0 ..— I — -—_ .- — — Tiine,seconds FI~UEE4.—Tfme hktory of a mfld pjwer-on pqE+Ip at 110mfles Wr IMIX ~-, (RmI.No. 7S) . G ~L2 C.G.occelero I 1A I I fion I I 1 1 1 I I I NI L 1 I 1 I I 1 I I RI , * , , , , I , -1 , , I m , ,.= , -, I I I I I I I I I 1 1 I 1 I 9 , “,. i I I I 0 I 1 I ., .5 ..., I rr,l 1 J&/ 1 7er wmg 11.0 - I 1 1 I I I I I I I 740.: I I I I I 2.0 1.5 -~me, SCCot7~s JIOmE #.-Jl’lme hfstom of m sbropt POW-4 Pnlkmat70mile p ,, -IDPQ<< 2doo , , I 1 1 1 1 1 I I — 1- & u -mu > ““ ~ hoar. @UU No.U L Time, seconds Frumu 4tL-’IYm6history ofan sbrnpt powrr.oE pn!lnp at 99III&Y pW ~. (Run No. 192) I 1 110, ..- 1 1 1 -.. , I t Ie voror < yoo ,+ < angle - ~. 1 1 I I 1 30 — 7 *“ \ 20 + Q. ? Q& Q190 ; 480 70 2 I I I I I I c;G. occefero+ion I , I Al , I I I 9 C? I I I I 1 1 I , I 1 I 1/I /1 1/ I 1A \l \A v I Y I - ! I RX. bcce~erafibn Tail I I I I I I I I I I I I/1 1 I , . 3!.0 $]~ I Vfl A % - I i.o I w I I h. I I I \ . I I I %..5 1600! I I II I I I l\l I -g >’.~ [ \ bch!y .’ “ Lci - “5$$$ m .fi+~*-l- ‘-s o --# i ii ii d s I ~., m IJ+M I I .5 I I I I 1.0 I I I I L6 1 1 1 I I I I I 2.0 I 1 (JX u ~ Me, seconds FIaum 47,-TIIIv3blat.my of an almmt Pow~*fl (Run No, 188) P~lWI at l% ~]- Pm ‘Wr 150 ~ --- I II , , I?l 1 I I -. I I I 1~ 1 1 I I I 30 .9evofo”r gng>e ‘t I i 1 I I “ I 20$ “oh-i-l-w ‘ ‘“me’ “h’”d’‘ ‘ ‘ ‘k>:. z-’”Ot--R--R u. I , , , ) , I 1/1 I , 4 1 1 I 1 I , I , w I -i- 1 ?ZZk?,Seconds Fxmma 4s.-Tlme Mstary of M ●brupt Powwd pull-up at IW mtlm par hour. (Run No. x9) --Time,seconds FIGUM 49.-TLme bbtory of sn abrupt power+ll “- pnU-uPat lW milespw hour. (Run No. 105) . -- Elauns MI,-Thna hi.dmy of an abrupt IIowerdf pull-up at181mll~ per hour. (Run No. MS) m El I — 726 REPORT NATIONAL ADVISORY CO-E FOR AERONAUTICS ~ 120 #6 c /00 Illllllll!l! I 1 1 I ,. }” -+Ak , speed ti t t C#Lw .- I I I 1 I 1, I J . 1: 1 ., ! ,-L I loll foaa -t, / 10 I 1A k — — ..... . .- ,....=____,= . u. s -400 / -: I “ A 1 k , 1 , , r , , , , , I , , Llhter’ I , , I , 1 , o I .5 I l.i I \ : /.0 i I 1 I 1“ I I 1.5 1. I I 2.0 1 1 . I I 2.5 Ofa mild ~WW-Jft pti+lp at 110miks w hour. (RUR No. 1$7) 1 {$ ~ Q&& L4 - 7Zme, seconds Fmurm 61.–Time hh~y ..JiF1 I \ _ I winal 1 1 !J I d I ~ -1- PRESSURE c DISTRIBUTION OF A PW–9 , I PURSUIT AIRPLANE 727 IN FLIGHT , Iml’ulmm -. — -. — -- “ 3m 0 -300 m Iil!iiiiiil — I I I I 42400 I lhl II r# 1 W[ f 1 I [\t I I I I f I I ~ ower 1 wi~ 1 I -k I .— .- I -4P= I m \ 1 hod-, r 9 t -. — I ~ Iiiiiii .5 0 -iii .5 I 1 i,li, 1.0 I I I l\l f I [ 1 ,,, ,,, l,l, 1.5 2.0 seconds Time, FIGURE5Z.-Tfme I&bxy of a mfld poweral p+III-upat 109rolhg E hozr. l,m<=l I u 2.5> —.— ~ (Erm M. 1$S) - —— 728 REPORT NATIONAL M) VISORY COMMITTEE FOR AERONAUTICS f lo 1 ; l.lmtlllll dl 1 I 80 70 :2 1 I I \ 1 I I I I I I 1 I I I I I I Y 1 I I , , I , w , , I I I VA \ I I I I t f x I I ~ 30$ $- I ! CG. ( g $+0 I , 1 ,Elevofa’r onoL.=._ I i 1 I I I I I + 1 I I I R’ [ I I I I I I i >W .\ f 1 I ..- .— j~\301 1 I I 1 I 1“ I ] Upper u .a , , 1 1 1 I I ..- 1 wing’ Hlltllll”t .- I.u Time,seconds FIGURES3.-Tfms hfstory of a mild powerdf ,= {.a .au c. -Lu puil-upat 107mfks psr hour. (Run No. lW) FJ .- 729 —— — .— Iiiii I I ! $3 $ n <2 * Q u t ! iii b 1 I I I I I I 1 I I I I I I I ~=- I I 1 . I 1 I . I I 1 I I 1 I 1 b I I 1 I !Tml .—. ck-cek.rnfirn —--— .—. —..- I t I I I 1 I I I 1I -i 1 I — I “ I I I I I I I .= —.— t 1 Ill I I Vefoc!”ly I L4fkulor. — —. — .= — : 1 I 1 I I !!!!!!!!!!!lwH!- $! [ IJ lilllll~ ;: I . I 1 I%7g occefe’rn)i& C.G. accefera+ion I I I — t I I “ \- . “ .— I ‘ .--— u. _- .50$$ Lw ~\ I .—— 2538 I :1 .. .— 40 m -- -6a .. I , s /m , , I /1 t- , , , I 1 [ 1 I I r I f 1 1 I 1 I 1 I ! ! 1 t t 1 / t Lower wing 1 1 t i A * , I I , t 1 —, 1 I I I 1 t I 1 I-Mm l-i+ — r 30 Illlwb I i I 1 111 1111 per wing t— — r t—— — — 50 60 -g 44 ~ # 40 m — ~ J- -.s FIGUEXs4-TimE 1.0 I [.5 2.0 Time,seconds 1 L 2s hktary of a mIId poweralt pult-np at E% mfks psr hmr. 3.0 (Run hi. ~ 3.5 .— 730 REPORT NATIONAL ADVISORY FOR AERO?ikETICS COMMITTEE ● # # I 1. 1 !-t, r , c 1I ardle -.I I I1 v , f 1 I 1 1 ! 1 1 I I, I , I, 1 1 I f . II I I 1 I 1 . -, 1 1 I t . \ Em 800 -60 400 0 -m moo -@o 4000 I I I I I I b!!be; 1“ I I I I I who Ichf.1. 1 1 a 1 1 I 1 1 1 # 1 8 7 1 Lower win; Iaod -, u 1 1 I 1 ! 1 I-I=L II I 2.G ‘OoEiEEh@ /.s 1- . ~ 1 1 !!l- ~nl -- 0 1 1 .5 1 1 1 1 1. . -, 1 I 1 1 1111111, !.0 .1.5 Time, seccmds FIGURES6.-TIme history of a mild ,,, 2.0 powemff poll-up at 140mllea per hour. ,,, 2.5 (Run No, Z@) .30 ‘1s0 ‘- 44 $ 3y3 ~ 4, z .$2 ~ I(qli /0 8 6 Feef a’ong 4 2 o“ o 2 4 semispan 6 feet fz 14 16 alon~semlsp% FICItIR~&S,-Span load OUCved. Reducd to same area showlngj the vnriatlon in rolat!ve dhtrlbuffon at dIUerent stagedof the maneuver. Pull-up d 70mllm per hour. (Run No. 661 66 55 44 ‘. 3’ ;3 22 (. Ii 10 B Feet 4 6 olong semiapon .? 00 2 4 6 a 10 fee+ a/on~ Semispok” 12 ““ /4 IWmm ti7,-Epanloadcnrvea.Rodumd w sama arcs ahowlngthe varlntlon h rdstiva dlatdbutlon at dlUe.rantstsgedof the marrouver, PuII.uP at 187mllm par hour, (Run No, 182) /6’ “- m o w PRESSURE DISTRIENJTIOX OF A PT–9 upper and lower wings at high angles of attack agrees remarIiably well with the load cumw deri-red from the design rules, stilI, the discreprmcy between the actuaI center of pr=ure and design center of pressure (3 I per cent) results in an unsafe condition for the front lift truss. Figure 60 shows the comparkon between the actuaI gross spar load curves from Run 132, and the design load curves (dead w~~ht not subtracted) for I@h angle of attack on the basis of the same load factor or totaI load. A smtdl difference in relative wing load esists between flight and design rules (1.37 and 1.29, respectively, for this case) which ma=~es the effect of center of pr=ure discrepancy on the upper wing to a alight extent. It is seen from the figure that the greater portion of the front uppersparis overloaded, but that it is underloaded at the tip. This latter condition is caused by the rearward d.ksplacement of the center of pressure at the tip sections, which is characteristic at high angles of attack. The effect of the actuaI load distribution on the primary bending moments is shown in F~e 61. These curves are gives for illustrative purposes only, but are true comparisons on the assumption that a pin joint exists in the spar at the cabane strut attachment poin~ whichis not actualIy the case. Theyshow,however, that a considerable discrepancy e.tits between the bending moments actually obtained and those assumed— in this particular case, 36 per cent on the unsafe side. In view of the importance of the l@h a@e of attack condition in design and of the position of the center of pressure in this condition, it would seem advisable to shift the design center of pressure forward on the upper wings of biplanea by the amount indicated for the combination by wind-tunnd tests o~ theoretical calculations, if such are available, or if not, to shift the upper wing center of pressure forward arbitrarily by 3 to 5 per cent to be on the safe side. It might also be advisable to increase the design bending moment of the front spar in the middle of the bay of externally braced types by an arbitrary amount to allow for the rearward displacement of the center of pressure at the tip, except in cases where the effect of tip shape is definitely known. The existing rules apply an arbitrary increase in bending moment of 30 per cent fmm the outer point of inflection i% the tip, which takes care of possible increases in stress near the strut attachment point arising fmm excessive tip Ioads. The tip loss assumed, which is considered to be more than that actually encountered, is supposed to take care of possible excesses of stress in the bay, but is not sutlicient to provide for conditions arking from displacements of the center of pressure from the assumed vahe. The above discussion accentuates the need for extensive research on load distribution over wing tips, which must be done before the arbitrary nature of such revisions as suggested can be eliminated. PURSUIT AIRPLAXE IX FLIGHT 7’33 Figures 62 and 63 are presented to show the variation of spar load distribution throughout two pull-ups of dit7erent character. Both show the same generaI resuks for similar portions of the maneuverl the outstanding points being the tip peaks on the upper rear spar in the region of l@h angle of attack and the si.darity between front and rear spar load distribution on the lower wing in all conditions. F~ges 30 to 55 represent histories of all of the pti-ups investigated for which satisfactory records were obtained. They show the relation existing between the loads acting at any instant during the maneuver. It is seen that the upper and lower wingg reach maximum loads at the same time; also, that the maximum taiI load is down and occurs early in themaneuver, well before the wing loads reach their ma.tium values. It wilI be noted, too, from the taiI acceleration records that this down load is not accompanied by an appreciable weight or inertia load, since the acceleration at the time is approximately zero. This suggests that a critical loading condition for the fusdage is the pure maximum down load on the tail, which agrees with the present design rules. Referring to the abrupt power on pu~-ups in the high angle of attack condition, the acceleration at the taiI remains at a fairly constant ratio to the acceleration at the center of gravity, and is of the order of one-half of the latter value (considering Ig the datum). It would fo~ow from this that the present rules are in error in assuming the l@h angle of attack condition to be equivalent to a static condition with the basic load multiplied by an appropriate load factor, and that they shouId take account of the fact that in this condition there is, in most cases, an angular acceleration about the pitching axis which radti in a varying inertia Ioad along the fusekige. This inertia load, of course, depends on the moment of inertia of the airpIane and the effectiveness of the controIs, and differs In any case, however, the with diBerent airplanes. condition would be such that the inertia load forward wouId be greater than anywhere else. This points to the conclusion that the present rule giw.s design loads for the engine mount and forward portion of the fusdage that are too sM, and hence m8y prove unsafe. Referring now to the histories of the power-off puUups, the above discussion does not apply. The tail accderations now are about the same as the center of gratity accelerations and usually a Iittle higher. In short, the conditions in the power+ff pull-ups are in fair agreement with the conditions assumed in the design rules; the conclusion is therefore drawn that two possible critical conditions exist for the fuselage at high incidence, one corresponding to power on and the other to power off. .— .— .— - -— . —.— . 10 , II 1 8 Fee f%onq II 4 semispan 2 -E 60 1,,11 2 4 6 fee+ o/on~semispu;” /2 Fmwm @J.-Cbmpwlson of high anglsofMa& spar loadshorn fllght W@ (rrin No. 1E2)and design rulfs -c -a & Q o ~uu 2 4 61.--Uampwlmn of tit 6sx _ @JJlt~al ,1 ..8 /2 10 Feet a/ong semispan mornem fmm tMght* at abmw atmamm$) (I= No. W /4 M sp.cddcati. 16 (PIu .. Lower UppeJ- wing wing Lower wing . _ ---i .-.111111111 I I I 1 100 0 1 .25 I r I 0 0 too 100 a 200 /00 100 ! o /00 ,50 200 L/qoerwing 1 L /00 .48 too 0 I 1 ---- 0 I I IJ --t W 00 75 0 z /,75 ,98 (Peek) 1,29 300 300 200 200 /00 /00 o o /08642 Fed Bmunx &2,-Thno history of spar obg # / - _ - - - -- --~...--- - 2468 10 la lea@ [n n pull-up at 187mllm por hour, (Run No, 132) 14 m /00- ———= = ‘ ——- . 00 100 275 span /00 -- – 0 Feet cdongspon lfmmix 6S.-Tlmo blstory of am bads In a PUU-UIIat llo mUW DW Mum (Run No. 7’2) a ’37 736 REPORT NATIONAL ADVISORT A point of considerable interest is disclosed by the results of the power on pull-ups. It has been mentioned previously in this report that one reason for investigating the pull-up at such length was to doterti’e the efFect of the pitching motion on the center of pressure position in the high angle of attack condition, and also on. the m8xirmum normal force coefficient. It has been shown that the maximum forward position of the center of pressure is the same within the experimental error, regardhss of the character of It can, therefore, be concluded that the pull-up. pitching dow not affect the C. P. in the high angIe of attack condition, at least for the airfoil section and cell used here. The maximum normal force coefficients are erratic; that is, for different runs the values of normal force coe~nt are not the same, and seem to bear no clear relationship with any other variable. This is believed to be due to the lack of accurate synchronization b~tween the air-speed record and pressure records. In general, however, the values are greater than would be expected. The upper wing CN from the model data (Reference 14), corrected for scale effect reaches a maximum value of about 1.43. The full-scrde results show some maximum values as high as 1.8, with the average maximum about 1.66. In only one case is the maximum CN less than 1.43. This would indicata strongly that the normal force coefficient of an airfoil with a positive pitching motion attains a higher value than a similar airfoil under This indication is further subst-ansteady conditions. tiated by Figure 64. It will be noted that the accelerations in the power on pull-ups lie close to the theoretical curve, NQw, since the theoretical curve is based on the =sumptions that the airplane is pulled up with no loss in &ir speed, and that ON ~,c~~ti~for the pitching airfoil is the same as that for steady fight at stalling speed, in view of the fac~ that the air speed actually does fall off an appreciable amount by the time the peak load in a pull-up is reached, it can only, be concluded that the close agreement between the experimental points and the theoretical curve is due to an unusually high ma&mm normal force coefficient occurring in the maneuver. This close agreement, therefore, could only be expected to occur for certain conditions in which the decrease in air speed just offsete the increase in normal force coefficient. In these tests the conditions are met in abrupt poweron pull-ups started from horizontal flight. In the power-off pull-ups the air speed falls off more rapidly, and as a result the measured accelerations in general come below the theoretical curve. It might. be expected from this that in abrupt pull-ups made from steep dives in which the weight component still acts forward at the peak load, the air speed would not fall off so rapidIy, and as a result accelerations in excess of the theoretical would be experienced. While none of the present data show this to be so, it has been demonstrated in maneuverability tests on both an F5’C% COMMITTEE FOR AEROX4UTICS and an F6C-4 airphme (results not yet published) that this condition actually occurs. These high normal force coefficients can probably be accounted for on the basis of tho known phenomenon of vortex or eddy formation. An important consideration @ these formations over a wing beginning to stall is that time is required for the back flow in the boundary layer to reach the stage where the &ret vortex can be ftied. In “the pitching wing, therefore, the flow — does not instantly break down when the steady fligh~ angle of rnmimum lift is reached, but continues in force for a short timo whale the wing rotates beyond this “tihgle, with the rwdt that the lift coatinucs to build _up to an abnormal value. Whatever the cguse, the result is interesting and important, as it can very conceivably have some eflect on high angle of attack load factors, and also on the relative distribution of load between upper and lower wings of a biplane. With respect to this latter point it would seem reasonable to expect that the lower wing normal force coefficient at peak tctal load in a pull-up would be nearly equaI to ita steady flight value, since the effec~ of the upper wing in the latter case is to delay the ~~eakdown in flow over the loiver wing, and henco the explanation given for the upper wing does not ““ apply, at least for the peak total load condition, Fur@rmore, the slope of the lower wihg CN curve is progressively decreasing with increasing anglo of attack (Reference 14), and at the time of maximum load on the upper wing has a fairly small value. This is borne out by the relatively ffat lower wing load curve in the present data. The above asstimption is substantiated by the facts, the model lower wing C~ at the angle of attack of maximum upper wing Ioad being 1.18 (corrected for scale effcct), whereas the average of the maximum full scale lower wing normal force coeflkients from tho power on pulI-ups is 1.23, a good agreement, If the hypothesis concerning the high upper wing normal force coe%icients is correct, the foregoing lower wing values can be made to agree still better by taking the modkl CN at a higher angle of attack. T@ effect of the phenomenon discussed in the above two paragraphs on the relative wing-load ratio would be, then, to increase the ratio for the pitching high angle of attack condition. Comparison of this ratio from the model data. and from the full scale pull-up data shows this to be true, the former VRIUOat maximum upper wing load being 1.16, whereas @ latter value (average of all power-on pull-ups) is 1,35, The ratio of these two values, since the lower wing normal force coefficients me essentially the same, is then practically equal to the ratio of the steady flight upper wing maximum normal force coefEcient to the pitching upper wing masimum normal force coefficient. In all pull-ups, the acceleration at the wing tip is of approximately the same magnitude as the acceleration” at the center of gravity, differences being a.c- PRESSURE DISTRIBUTION OF A F’iV-9 PURSUIT counted for as foIlows: in the edy stages of the maneuver, the flexibility and lightness of the wing structure aIIow the wing tip to accelerate more rapidly than the fuselage; in later stages of the maneuver, slight ding motion is undoubtedly the cause. .4 phenomenon ewkk.need on the original records, but not given in the histories, shows the esistence of a high-frequency vibration of large amplitude at the wing tip occm-ring in each case at the iustant of mtimum load and continuing until the end of the record. (F%. 65.) Whether this oscillation is caused by the sudden breakdown of air flow on the wing tip, by the abrupt change of inertia load, or is associated with the more or less welI-knowm phenomenon of wing flutter, Vibration tests conducted on can only be surmised. the airplane showed the natnraI frequency of the wing structure as a whole to be: (a) From S.7 to 10 oscillations per second in bending, the differences being caused by the addition of various weights up to 22 pounds at the accelerometer location; (b) 8.5 oscillations per second in torsion. For tho upper tip alone, a rigid support being provided at tie strut attachment, the frequencies were taken from 10 to 14 in bending with weights added as before, and 12.5 m In the bending tests, the higher frequencies torsion. corr=pond to the condition of no estra weight, which is the condition occurring in flight. It was diflicult to determine frequency of oscillation from the wing accderometer records because of the superposition of several waves of dif%rent frequencies and amplitudes, but in general the frequency of the principal oscillations was of the order of from 10 to 17. Strangely enough, the higher frequencies occurred in the slower power-off maneuvers, thus precluding the possibility that engine tibration was responsible. RoIIs.-Figures 66a to 67i show the distribution of Ioad throughout a right and a left barrel roll The initial stages of the maneuver up to the peak load are identical with the pull-up, ~~cept that the loads build up on the rudder in addition. Beyond the peak load, autorotation starts and is evidenced in the right roI1 by the shapes of the pressure curves on the right upper wing which are characteri..tic at angles of attack above the stall. (See Reference 14.] The left lower wing in the right rolI does not stall, the condition being similar to that at high angksnearnmximudift. In the left roll, the right upper wing is not stalIed, but is near or at ma.sirnum lift, while the left lower wing shows etidence of being st.aUed very slightly. The horizontal tail surfaces show disseymmetqy of load. In tie left roll, the distribution of pressure on the right side is very simiIar throughout to the pressure distribution in a pull-up, but in the right roll the down load on the elevator is replaced by an up Ioad in the latter stage of the maneuver. This dissymmetry is probably caused by the influence of the fuselage. .41RPL4X~ IN FLIGIIT 737 Span Ioad curves for the two roIIs are given in F@res 65 and 69, and time histories in Figures 70 and 71. ~ Figure 72 is a span load and inertia load diagram combined from correspon~u points in the right and ; left rolk at the condition of masimum dissymmetry of load on the upper wing. This point occurs at. 1 66 and 67 that { second. It mill be noted from F~es the two maneuvers are closely similar, viz, the control ~ action is practically the samel and the rua.simum ~ accelerations at the C. G. are practically equal and occur at the same time rdative to the start of the maneuver. Figure 72 is corrected for the difkwnce in total load in the two maneuvers. The outstanding points in comection with the unsymmetrical condition are: 1. The masinmm dissymmetry of load on the upper wing occurs shortly after the masimum total load, and the totaI load in the UUSPIUM trical condition is not much less than this maximum. 2. The diss.pune~ on the lo~er wings is of such a nature as to oppose the rolling moment due to the upper wing. The first point is of particular interest in that. it shows that the unsymmetrical cabane load condition shonId be analyzed with an average load factor equal to the high incidence load factor. This, of course, applies only if the arbitrary nature of the &sting rules. is eliminated, and they are reformulated to include the effect of varying inertia load across the span and the true effect of the lower wing. The second point is of considerable interest &o, altbough in the light of the delayed stall of lower wings of a biplane indicated by model tests, the resuh is not particularly surprising. (Reference 14. ) The remdts indicate that at the point of masirnum upper wing rolling rno~ent, the angle of attack of the strolled side of the ce~ule is not beyond the st~~ angle of the lower tig on that side. This would explain the count--rolling moment of the Iower wing. The inertia load along the span varies as shown in the figure. The slope of the line is calculated from the angular acceleration of the airplane (5 radians per second per second) as determined from the slope of the angular velocity record for the right roll. The wing tip acceleration for the right roll checks this slope very cIosely, although the wing tip ~cceleration for the left roll does not. It maybe of interest to note that the angular acceL eration as calcdated from the known rolling moments and moment of inertia of the airplane checks the -due from the anguIar velocity recorder and accekrometer faidy closely. This can onIy be done, however, if the total wing rolling moment, as determined from F~e 72, be assumed as that for the right roU. The reason for this is that. the anguhw -velocities in right and left rolls are not. the same on account of propeIIer torque and because the load on the stalled wing in a roll is quite sensitive to changes in rolling -relocity, whereas — -. .— ..— — - — ..— — —. . —— .-+- .— /0 /0 — 8 I 1 1 I ,. m x I 0 Thrn&ica/(V = 60) ulflu,s. Mwer-on X Power-o 47 x # pull-ups. 1 1 i“ t & 4?6 ~ ~ b .* 4 0 x + d~ 2 c/ . _ o {00% 20 < / / / 1 1, >y H . ‘ / 40 FNXJBE6&-Thwre@l 60 120 80 /00 Inifiul oir speed, m.p.h. and erperlmenti 140 160 180 Tree, secon@ FIQ,um66.-FWpmduotkm of wing-tlp amshrom- rwmd (run No. M), &mrao@r of vibrstlon Mter peak loaU SPPIM load faotarsin abrupt pnll-n~ from level Uht FMUBEm I ,, thwing Frau’m em mlu’mMO a h. . mamw 6cd Bmntm 00f ● ,: “fl FIOW C4h . mum :, Lw — -. .,. ..—---- ... .._ ,———.- — -- -—.. .. . ——————.—— lv— Fmunc 07a Run 225: Lcff barrel roll o+ f63m,p, h. Time;. 75 sec. v/ ~10UB6 . t17d .n . FIGUBB67f FIGUIU67e ,,, i’. i! I ?mum 67g mom m feet olonti semiapan ‘1 b-r-l .-!::- I ..1 I .r- I I +01 Fli JO a I I -i-t 4“”2”-Ou 6 Fee+ along molmx I-t .-. -.— I -. -. L& semispan FIaurm @8,-V@rlatIrm of span load dfatrlbutlon in a right borral roll at 187mllw per hour, (Rnn 1171 200 c & 100 I .- fo I I I ‘..”0sec. 8 Feet %ongsem?spon 2 I Wln r,oof~n, 0-0 Franx W!.-Vnrtatlonof~ ,!, I ,’ ,11,1!:’ 1 ~fiiijiiiiii~lp,,lll sec.” 2 4 6 feet 8 along I I I 12 semispa~” lord dfslrlbulIonIn u loftImrmlmll ●t la mlh B hoar, (Run No, 2!M) 1 I I I I 1 /4 1 1 I 7 16 No, 222) 744 REPORT NATIONAL ADVISORY COMMITHZE FOE AERO.SAUTICS ... -. —. .W+H+tH+ — “+IEEERL ,. .. ,U [4 I II II I . .. — t ,- tAr@or’ I I I I I I 1 I I I ve’bcif I 1’ , [ ~ , ,,. I t“ I .- Y foif Iood. Horizontal 400 g? ! !I I 1“1 I !’ +- 4 : .- —. -400 I r 1 1 1 6 1 , 1 4 .- /7 I I I I J T I =l!L I i gti .> -5000 I ~ , I Abrizon+ol I I t I I I I I I I .1 I I toil momen+” I f I I I l:! I I l!.!ppe< . Wingl 1 l— I ! I I I 1. I 1 I [ I I !. I 1 Im I ~1 I I I I 1 1 I I I I I 1 I 1 I I , I , I J I ~1 1 I I I I 1.1 I I I 1 I I b Lower wing., I I I I 1 .. —. — .7Lme, seconds FIOURE70.-Time bktorg of a @t bmrel roll at 167miles w how. (Run No. 222) - PRESSURE DISTRIBUI’ION OF A PTF-9 PURSUIT AIRPLA3ii 745 LW FLIGHT * — — . .— t $ \o — .= — . .— — b .. I I I I t I lTud c&ce/erafion ~ _= u .— — -,..- . — — I 1 I 1A t 1 . I I I 1+1 l-i-k I t 1 t I I $1 I 4aoo3 ..-— -. — ! I Lowe ~ 64H4H+H . .. -.— _= .- ---— -= —. . - -. ._ = -. —._ —. .= Time, second.. ● 746 REPORT NATIONAL ADVISORY the load on the unstalled wing is not. Therefore, if the right roll is chosm, we have the actual measured value of the load on the stalled or right upper wing, whereas if the left rolI were chosen we should have to assume that the load on the right or stalled wing in the right roll was the same as the lotid on the left wing in a left roll, which would be a rather hazardous assumption, Assuming, therefore, that the wing-rolling moments given in Figure 72 represent fairly welI the true wing-rolling mornant in a right roll, it is onIy necessary to subtraot the opposing torque of the propeller (in this case approximately 900 pound-feet, based on a knowledge of the propeller and engine characteristics and the conditions at which they me operating) to obtain the total moment causing angular m~>terd CO~ITTEE FOR AERONAUTICS they assume a major importance in the study of the phenomenon of autorotation. Figures 75a to 75q and 76a to 76p show the distribution of pressure for a right the and a left spin. In the right spin throughout, left lower wing remains essentially unstalled, but at an angle very close to that of maximum lift, whh the right upper mung begins to show evidences of becoming stalled at the tip at 2 seconds, imd thereafter the stall progresses from the tip toward the center until at 2% seconds the entire wing is stalled to the plane of symmetry. This condition continues, with the load becodg smaller, to 4 seconds, where the loads start to build up rapidly and the stall apperm to become more pronounced; that is, the center of pressure locus moves back, indicating an increase in angle of attack. L?> 7.42 fi i%eef afonq spon Fi@JWE7G-Span I&IdOU?VH fromrfghtand left rolls mmbfnd to show worstunsymmetricalcondItlon acceleration. pound-feet. This value is: 9,600-900, or 8,700 M a, therefore, is — I or 8,700 ~ = 6.13 radians per second per second, 2 Time histories of ~mht and left rolls at slower speeds are given in Figures 73 and 74. These runs do not fit together as well as the higher speed rolls, but it is apparent from a study of the time historithat the time of maximum dissymmetry occurs relatively late in the maneuver with the total load from 80 to 90. per cent of its maximum. Spins,—Loads in the spins are relatively small and uninteresting from the structural point of view, but The wholo motion appeam to be unsteady up until the records were discontinued, although it is quite possible that a steady condition would have been reached had the spin been continued longer. Loads on the hori---zont al tail surfaces remain characteristic of high angles of attack with the elevator well up until rtbout 5 seconds, when the load builds up in the po”sitive direction, the pressure distribution becomes more irregular, and the eIevator load changes from negative to positive. Thesq changm occur ;Vithout movement of the elevator and indicate a change in ah-flow conditions, probably caused by the change in direction induced by the increased anguhw velocity. It is interesting to noto from the tie history (fig. 79) of this maneuver in conjunction with the pressure distribution curves that the horizontal tail moment changes early from positive to negative and continues to build up in tho negative or ‘ PHESSURE DISTRIBUTION OF A PW-9 diving sense to a conaiderable value, whiIe the airplane is still settling into a stronger spin. Even the load on the elevator changes in direction from down to up with the angular displacement remaining constant at about 24° up. In the left spin (figs. 76, 78, and 80) tho right upper wing remains unstalled throughout, but the left lovmr wing appears to be stalIed from the beginning of the record and continues so until the end of the record, the stall becoming more pronounced as the spin conThe motion, as in the right spin, is unsteady. tinues. As in the right spin, also, the horizontal tail load increases upward and the trd moment continues to build up in the diving sense, while the elevator remains we~ up. The distribution of pressure, however, is more regular than in the right spin, undoubtedly because the right- side is no-iv advancing into the wind with the fuselage and vertical tail surfac% behind it exerting little disturbing influence. It wiU be noticed, too, that the horizontal taii moment in the left spin does not reach as high a vahe as in the right spin. ~ertical tail moments in both spins remain nefdy constant and at about the same value in both. It may be lemarlmd here that this particular airpIane was easily controlled in the right spin, but came out of the left spin with considerable difficulty. lmerted flight.-Attempts were made to obtain records in the inverted flight condition without much success, as the pilot found it difEcult to maintain the condition. The inverted attitude was roached by means of a half loop, but could not be heId, the airplane losing altitude and having a strong tendency to nose down and continue the loop. Figure 81 shows the distribution of pressure in the inverted condition. It will be noted thht the center of preEsure locus is rather far forward, indicating that the angle of attack has not reached the mgle of maximum negative Iift. Span-load curves me given in Figure 82. Dives,-A representative dive is ilktrated in Figures 83. to 86. Leading-edge pressures reach an exceedingly high vcdue, as noted from Figure 83 and Table III. The span-load curvee (fig. 84) show the tiect of the twist of the upper wing, the load inboard being positive with the tip operating below zero lift. A curve showing the variation of moment about the leuding edge along the span is given in Figure 85. The spar-load curves given in F~re 86 show the efkot of the washed-in centraI portion of the wing very clearly and also, to some extent, the effect of the washout, the front spar load increasing toward the tip with the rear spar load decre~~ing. h’ormally in a dive, m in the caae reported here, zero lift is not reached, although it could be attained if the airphme were rosed over sufficiently. This condition, however, is an uncomfortable one for the pilot, the sensation being that the airplane is slightly over on its back, which it really wotid be in a majority of PUTMUIT &EtPLANE IN FLIGHT 747 cases. Zero lift can be attained, however, with the nose of the airphme only dightly beIow the horizontal in a sudden push into a dive. Pressure distribution for this case is given ia Figure 87; it is sirnihr to that in the dive, but with the negative areas at the leading edge in greater proportion. The pressures, however, are much 1sss because of the lower speed. Span-loading curves in F~re 88 show that a considerable portion of the upper wing is at negative Lift, while loads on the central portion are positive, and that the lower wing load is positive throughout the span. Spar-load curves me given in Figure 89. The effect of the twisted upper wing is apparent in these curves, the down load increass on the front spar toward the tip and the up load on the rear spar decreasing The condition woukl be much more severe in the case of a fast dive at zero Iift., since the twisting moment wouId increase as the square of the speed, and hence the deformation would increase, remltiug in a greater proportion of load on the outboard portion of the hont spar. It is conceivable that the outboard rear spar load might reduce to zero, the entire load being carried on the front spar, in which case a form of partial inverted flight condition would exist that would probably be critical for the Ianding wirea and the leading edge of the wing. Pull out of dive.-An interwting, though isolated, case of a low incidence loading condition occurs in the pull out of the dive, RurI 226. h connection with this condition it should be said that the puIl out was normally executed; that is, it was made cautiously with due regard to the speed at which the airplane was traveling, and that it in nowise represents a speciaI t-t condition. The point chosen in working the records W= the point of rmdrnum acceleration (3.69), which occurred early in t-he pull out and which was felt would probably represent the most severe rear sparIoading condition that could be found without working up a Iarge number of points.. Figures 90 and 91 represent the distribution of load for this case. The condition is seen to be characterized by moderat.dy high load with the center of pressure fairly well back and span load tapering off rapidly toward the tip. The spar loads given in Figure 92 show the tiect of these characteristics. Spar loads derived from the speciiied Iow angle of attack loading condition are given for comparison, although the total loads for the two cases are not the same. The total load for the observed results is the true total for this case corresponding to an applied load factor of 3.6, whereas the derived curves are based on a load factor of 325 which is the low angle of attack load factor divided by the intended factor of sdety of 2. It can be seen that the upper rem spar load is welI with the design load, and although the front spar load is in excess of the specified load for low angle of attack, the condition is not critical for the member. On the lower wing, however, both sDars are overloaded. the front one afmin being -. .-— —..= —— ...— ,— .— -,— .— .- .- — — .— ..— .— . — _.— — 748 REPORT NATIONAL ADVKSORY COMMITTEE FOR AERON.4UTICS . , I , 1 .. I 2 } o Immm-lw , I , & I . ! $j low x4 .- g’ -. ‘A .0 — — he, $ekcmds Fm.mE 7s.-T[DM hfstoryof a fight barrel roll at 128idea per hour. (Run No. 220) PRESSURE DISTRIBUTION OF A PW-9 PURSUIT AIRPLANE 749 IS FLLGHT — .1 $ 1 I I I =. [ 1 I I , , “ , t 1 , , t 1 I , , , —HsH-H+H-tl’O . Ll-= i? ..— —— .—— — ---- . I d I I l\l RU I I I I I I t I b, t 1 r — — .— 1 1 1 1 t 1 1 I . — — — . I —+++5H+++-+-J =- E--I . uUso : .. I I I 1 I t I I t U.b&f+’@l--P”l I I I t 0 I .5 I .— — .-— - I I Id [.5 Free, seca-h4s 2d 2s FIGrim 7L—Time history of a Ieft bard mU at l!M5 miles per horu. (Run No. Z9) —. -— ,.. Fmm n 7Ea A I?launx Ii 7SC FI(fn’m 76d mmmn 7W nnmm 76! n FmuEE 75] FIoum mk ., . . ,, ., ~’ Fmwst W ,, ‘i’iui’l : Iil FlaunE , ,“ ,1 , Ii i’i’ I i, ,, ‘IISP Fmmm FIGURE7@ I FIGURE ~b 76a FIOUEX760 ,, 1 I ., ! . . —.-—.. -----. n -- ~- .. -—. . ———. .“. —.. ---- — n FIaum . , Iflouu 7of . ,’ 1 7ea ..... -.>, . --%: -~ -+%.. . . “’.. .,.. $ ..<.+ ‘s ‘.. <-:.., % I r ‘u lhamm 761 FIOVBE7m “1 ‘1 ,. . Run 227: Left sp!ti ?Zme: 475 see, y/ :— II I 160 /6(9 {40 i40 A . bwe} wing I I I 1 I ~ i.@er . \ .r- f+ .- . wihg I II .- r ~~sj _ - .- 1 au \ _ L . —_ : =- ““-I 10 I 4ot+ 40 20 20 I R I S fee+ along g:?: 4 semkpun ~ 0 2 Fmunq~.-Variation ... 6 Feef of aSWIlcmddiaiaibutlonIn a rightL@. ,, .— fed 4 2 ,,, 1 , /0 8 olong semispon (RunNo.:~) ,, ,, Fee} along oibng semispon RuuaE 7&-VarMbn & ma W dklbntlon h a m SPIII. (Run No. 2279 . seinispon 12 /4 16 PRESSURE DLS~UTION OF A PW-fJ PURSUIT AD3PLAWE 759 IN FLIGHT .. :,= -. .— — .. . — ~ H-H---l I I I I I 1. I 1 I I 1 ..= —. .— ,.= .- _ — -. —. I , 1 - 1 k IOCWQ d 1 + -. .-—= . .,,= —_ . —— == :— ... -.= _-- 200( — . 3 k v t I I 1 1 t 1 1 n’ ~,; I I I 1 II , 1 I.Lower wha t I I .— — . hod I ! ! ! Ower - 1 I I — Wr”m-. .= -.— .— .-.—-— ——-. = .— --Time, seconds FIGCEE79.-Time hMocY ot a @t sptn. (RUO h-o. ~ . . 760 REPORT NATIONAL ADVISORY COMWTI’EE FOR .— AERONAUTICS ● I , , t I , , I , L L , , I I 1 , , I .. I . ,. , t 1 I I I — ,. r 200 1 1 1 , , , 1 rl , . ..- 0 1+ 1 .iooo — -2000 .- -s000 1 , , , L A , 1 ___ ,-. Iiiiiil”ii” I 40 I 1 1 1 I 1 !,. K t I I II* I I I I — 1 .gg $ liiiiii , , r . . I -!i”kITTRQ_H71 Time, seconds FIGUEEEU.-The Meteiy of a left L?@, (Run No. M) -- r — Irmvm m BIOUBE 81 FIQVBE8Z,+pan 10M!tllstrlbutlon In Inverk%lfllgbt at 7?3mfled per hour. (Run No, 211) semispan ,Feef along FIOVBX8s,--Epan moment distrkmthm h a dive at 280rdh per hour. (Run No. !2MI.) (Momanta refamedta landlngedge at * pointad *) m , r Ill wer whg h 1 x ? — — I I 10 I trm~ FIOUBE 87 , 7nn h 150 ,— .. I 1 I I -m -mn I I I 8 I I I Feet%ng I ,I II I II # 4 semispon I I I 1 I I 2 1 K(M I ‘“” , ;+50 FItiu’m 80.+ ‘“”k hadhg in aIWOat M mih pm boor. (Mm No. 220) - :1 +100 +50 o -50 -/00 I Fee+ o/ong semispun [email protected] lmdln~aM*~lft[(uush-dow:at:lW:mllm-mei.houfi~m~No:.fl5) 500 ? ~417Q %00 $200 $ ~ 100 + (2 Fmumx‘JI.-tlpan lend dbtrlbution In a pull-out from d!va, (Ftnn No. 22S) Feef lhouRx w maurm M.+omparhn crhig orlow amh of awkl.mm semispan Ids fmm flkbt ~ (rm No. W and II* ,, ‘Uh 764 REPORT NATIONAL ADVISORY taken care of in the high angle of attack condition. On both win~ the center of pressure is somewhat farther forward than is assumed in the design, the total load is higher, and the reIative efficiency of the uppm wing is only 1.02 as against the design twsumption of 1.29. Of these factors, the first tends to decrease the severity of the condition as compared to the spcicifications, while the second and third tend to increase it. It is impossible to say from one case, of course, what center of pressure position is most Iikely to be encountered in the critical condition for the rem spars, since the load factor changes with the pressuro distribution, Defiite conclusions on this point, therefore, are not warranted. The comparison is little changed if both total loads are assumed to be the same, and the effect of the discrepancy betwem the reIative upper wing efficiencies remains as the most significant factor. The case, however, emphasizes the importance of extended research into the mutuaI interference of biplane wings, and also points out that the supposed factor of safety of 2 is not unlildy to be overworked, even under conditions which are not considered in any respect abnormaI. From the standpoint of the operating personnel, it wouId seem that the airplane should be handled very gently in pull-outs from fast dives untd the low angle of attack condition is more thoroughly known and provided for. Tail loads.—The important tail loade are summarized in Table IV. For each run ‘listed, several conditions are given in chronological order to cover the following: 1. Maximum horizontal surface down load; 2. Maximum horizontal surface up load; 3. Maximum stabilizer pressure; 4. Maiimum ele-rater pressure; 5. Maxifin pressure; mum vertical surface load; 6. Masimum 7. Wmimum rudder pressure. Vertical surface pressuras were not worked up for the pull-ups since they are relatively low and of little interest, Tail surface specifications for pursuit type airplanes impose average loads of 45. and 40 pounds per square foot on the horimntal and vertical surfaces, respectively. It is interesting, as a primary comparison between obmrved and specified loads, to note the values given in the columns of average loads in Table V. It must be borne in mind while doing this that the specifications are supposed to incorporate a factor of safety of 2; thus, any value given in the table exceeding one-half the specified value is to be considered. an overload and vice versa. It will be noted that the horizontal tail surface loads in the pull-ups are generally lower than one-haIf the design load, exceeding this value in Runs 134 and 137, and cIosely approaching it in Run 133. It is worth noting that the applied lorid factor (C. G. acceleration) in Run 133 is 6, making the safety factor for both wings and tail surfaces approximately 2 (on the basis of loads only) in the same maneuver, In the two dives listed, however, the tafl loading is excessive, being 26.4--and 30,7 pounds per square foot for Runs 213 and 226 respectively, CO.WITT’EE FOR AXRONATJTICS indicating that the spetications should be revised upward. Other high loadings on the horizontal surfaces occur in the high-speed barrel rolls, Runs 222 and 225, the values exceeding one-half the design loading, but remaining less than the dive loadings. It is interesting to note that in the rolls the max+mum up 10ad is of the same order of magnitude as the down .._ loads, whereas in all other maneuvers the down loads only are severe. On the vertical’ surfaces only two loads listed excccd the safe value, in the right barrel roll, Run 222, and in the pull out of a dive, Run 226. Since the barrel roll under discussion was quite severe and may be considered an unusual or test case, it would not be reasonable to e.spect that under normal conditions vertical tail surface loads would be so high. The load in the pull out must be co~sidered a fair one, however, and the possibility of the vertical surfaces receiving high loads under normal conditions can not be denied. In this caso the maximum load occurs simultaneously with the low incidence condition discussed previously in the report. Besides the average loading likely to be encountered on td surfaces, the specifications must anticipate the distribution of load with particular respect to the high concentrations that may occur in some places, usually the leading edges of stabilizer and fin. The present specifications dispose the loa,d uniformly o~’er the fixed surfaces whence itiecreases untiI, at the trailing edge of the ‘movable element, its value is one-third the wdue over the fixed surface. In addition, special leadingedge loads are applied extending from the leading edge back one-fifth the chord of the tied surface, and having a uniform vahle equaI ta three time9 the specified Thus, the leading-edge load for a average loading. pursuit airplane stabilizer would be 135 pounds per square foot and for the fin 120 pounds per square foot. On the horizontal surfaces the maximum pressures on the leading edge occur in the severe pall-ups and exceed the specified leading edge value by a very appreciable marginl although fortunately they are usually quite local in character and do not extend over the specified area. The fact that they not only exceed half the specified loading but actually exceed the whole of it is well worth noting. In all of these cases (pullups), however, the accderations at the C. G. wero In the less severe power+n puILups, greater than 6. the masimum prewres recorded are under] 35 pounds per square foot, but in some cases considerably greater than half that value, indicating that the specifications should be revised upward. In one dive, RurI 220, the pressurewas equal to. that specified within the experimental error, md it would seem from this that to double the present leading edge specifications would give a much more reasonable value. In no case does the pressure on the leading edge of the fin exceed the specified value, although in the PUU __ out it reaches a value of 90 pounds per square foot, . -. ._ — .... ~.. ._ ,,... PRESSURE DISTRIBUTION OF A PTV+ A few comparisons of some of the worst rib loads with the specifications are gi~en in Fi=~es 93 and 94. They are self-explanatory and need no further comment. Slip stream investigation.-A number of level flight runs and pull-ups, both power on and power off, were made with oriflw located on sis ribs in the centd portion of the upper wing, and one rib Dear the root of each lower wing to determine the effect of the slip stream on the pressure distribution. Referring to the leveI fIight pressure plots (figs. 95a to 95c), no pronounced ditlermce between right and left sides can be observed in the slow-peed condition on the upper wing, although an appreciable difference on the Iower wing is apparent, the effect of the rotation of the slip stream being, as wouId be expected, to reduce the eflective angle of attack on the right side. This effect on the lower wing occurs thoughout the speed range. On the upper wing at the higher speeds, a similar effect is noticeable, aIthough the differences are not greater than might be expected es a result of the deformed leading edge. The pressure plots for the peak loads in pull-ups show no appreciable differences in the character of the pressure distribution, with the ~~ception that the leading-edge pressure on the right lower wing is lower than that of the left lower wing in the power-on pulhps, whereas botk pressures are about equal in the power-off pull-ups. The span load curves show more clearly what differences exist. In Figures 99 and 100 it is seen that the slip stream increases the load on the central portion of the upper wing, while the shape of the load curve remains much the same. The effect of the rotation of the slip stream on the lower wing, however, is prcmmnced as the figures show. Figures 101 and 102 show that no appreciable dissymmetry of load exists on either the upper or lower wings at the peak loade of the puU-ups as a result of the slip stream rotation, aIthough in the power-on pull-ups the total load in the region fiected is greater than the load in the power-off pulhups for the same initial air speed. This increment of load is due to an increased air speed, which is composed of both the sIip stream increment and the natural increase obtaining as a result of the power-on conditions. From Figure 103, which represents the condition in a dive at high speed, it is apparent from the lower wing rib loads that the rotation of the slip stream is still efkctive. The low point at rib A’, which is also apparent in the other span load curv=, is probably not due to slip stream tiect nor to any interference from the fuselage. The pressure records for this rib show violent fluctuations of pressure for the points aft of about the quarter chord point. This fact, in combina- PURSUIT MRPUXE IW FLIGHT 765 tion ~th the knowledge that a rather abrupt discontinuity existed in the upper surface wing curve near the leading edge of fib A’j would lead one to believe that the streamline flow over this rib was disrupted, and hence caused a marked decrease in the Mt. This is further apparent from the pressure plots for the dive. @ii. 98.) The conditicm may thus be considered as purely locaI and having no connection with the slip stream. Fabric pressures.-C!oincident with the slip stream measurements, records were also obtained of the fabric pressurw on rib C and several other poin@. For this purpose each pressure capsule was connected to one of the flush ofices in the wing surface and to an open tube terminating inside the wing near the flush oriilce. Pressures as obtained me tabulated m Table V and need no speciaI comment other than that minus signs indicate downward acting pmsuree and vice versa, regardkss of whether the ofice is on the upper or lower surface of the wing. -= .— .-—— .--.— ..— —— CONCLUSIONS It is concluded from these tests that: 1. In any condition of flight characterized by low angles of attack with thg center of pressure well back of the elastic ask, the wing, unless quite rigid, may deflect torsionally to such an extent as to greatly aher the load distribution. This means that the load distribution is a function of the torsional rigidity of the wing structure and this fact should be taken into account in the design rules. 2. Regardless of the cause, the effect of wing twist on the load distribution is the same and this effect can be satisfactorily calculated. 3. The maximum forward position of the center of pressure on upper and Iower wings is unatkcted by the factors involved in accelerated ~Uht.. 4. The masimum forward position of the center of pressure on the upper wing of the fulkxde airplane is the same as that for the model wing. This point h questionable with respect to the lower wing, since strictIy comparable data are not available. 5. The mfium normal force coefficient of the upper wing of a biphme reaches an appreciably higher value during a maneuver involving positive angular velocity in pitch than it does in steady flight, while lower wing normaI force coei3icient9 corresponding to the mgle of attack at which the upper wing masimums occur are essentially the same for both pitching and steady flight. This means that the relative wingIoading ratio for the true high angle of attack condition (-which always involves a positive angular velocity in pitch) is greater than would be expected from steady@ht (wind tunnel) data. On the PIT”, this increase is about 16.3 per cent of the steady-flight value. ..— _.— — .— .— ..-C .— .- .—_= .— --— .— ..— --— .— — 766 REPORT NATIONAL ADVLSORY COMMITTEE 60 ‘“ 40 FOR AERONAUTICS /80 -— 1 Vi ‘ ~ 20 < al- ~ L.E. L +. –Minge$ /40. [ - - .= /20 - 1 k ---sp ; 100 1 , ; d leu di”)g d gelo 00’ ..— .— ecifie I -40 80 1 --LO o~ on rib fime 1 ’60 ,[ & LOQd -80 60- — $d 1 : in / 1 Run )%x rib 1 1 N — I%.[37, ?f3 40 + -MO 1 1 1 520. ~ rib N in Run NO.226 1 f{m.s 16°~5sec. 1 I 1 +- -+&ified leudlhg ~dge load -HO 1 1 1 sec. I 1 -1oo ~ ~--bad N in Run No. 13? / sec. ‘t –Hfhge @ T.E. \ o ~E ‘ . . ; ~ .2(J ‘--J : 1 I + g - \ \ .&. -160 / / -40 -/800 olong I d ’600 3 Fe& [ Feef olonq .. ch~ro’ 2 cAord 3 ~QURE W,-Comparison of worst horbmnt.nimrfaw rfh loads with spedtlcat[ons 180 ., Load I% rib’ s h . . .._ /60 Qii k 20 140 /20 ---< /00 /“ L.E. : I q-l 60 --- 1 1 f ! 3 Feef alw2g chord %: , ~qud”on rib S in Run No. 226, fime 23 ~ sec. ,, ~“. A%ge ... ~- .. . ~ -. ,-—. i (g-* 1 i t 1 , 40 / $ ~2 Ieud%gedgeloud t I 1 ~. ~d - Sped —. / ? --— . ......—. . ..— .— / 4 J 20 ; I I , 0 1 LE. “-— 3 Feef olo;g chord r2 FIGURE94.-Comparlwn of wod vertkd mrfaca rlb Io8dswith .?@tkations . — -. ,—..— ———. .... .--- -——... .—— —.—- ---—- ,. —-,. .. 1%1 r /+’,, \~ ---,—— ..--— ..-—. -— -. ‘-’)---—.-. ./’~~ /“” “-v ‘~’ . . > ,. ‘“ /’” ‘L.. - .,4 .3 /4 *f , ,7’ ““’%, ., ‘., ‘ “, B ,. . 80 ~% 40 g{ “ 20 $4 o?/ _&l Q. Run 302.’Levelflight d /14,5 m.p.h. — .——..- .. . ~x w !4 mum ml FIOURR ma I 1 -.y.. Fmu’M Mb L FWJBB ma mom 97% I ——— -—— ,..— - ,---- - --— ---- . ..— —— —— / - l—l—l—k - 80 I I I I I I I II I -k. I 60 40 c 20 ~ l— maum970 ~ P OB, % AM 4 2. .$ 2 40 20 n “ . H “ Fraum DD,-Spsnhmddlatrhul[on In dlp strewo L, — — . ..- .—. . mou’m w .. r —.— B A’ A A“ 4 1da 77U REPORT NATIONAL ~VISORY COlJ3~EE FOR AERONAUTICS — 2/0 180 t ,--- 150 /20 400 ‘ I 1 I 1 1 1 1 I 1 ! 1 1 I Is! {,i I Upper wing I 1 I I [ III I H -4. , .— 1./: I ~ .. I i, ‘ :?4 90 60 1 I t 30 ? * > ‘8’ II “AN R:n:2W A- sfeady ,, ~ S, Limifs’ ~ , & Feeiolo~gspun flig-hf of149m.p. h.pbwer ,, “Af9.4:*;ow:; ofprojecfedprope fil I q> ‘B’ 4 2 4 . .? ~ ~ I 111”1 A’” A,, 2 & 74 $ o II 1,/1 I A B A’ 2 4 on fler di k on wings Q/20 o &go 60 30 0 h’ H) c 0 2 4 Feef 2 4 $’ Feet o/ong 2 ... olong spon [email protected] load distribution in SI!P strum . 4 spon H“ 2 4 Feeialongspon 3GUIiE 101.-2pa?I md dbtribution in slip strsw.u. Pesk bgdr In power+m pcdhnps, feet olongspan $ ● Run323 of I.?lm.p.h. Run 322, of 75m.p.h. u 324, ,$M9# R “ *’ +“34” G5”” U x S, .Limifs of pro]ecfedpmpeller disk on wings 0 J ~ 400 k timo 200 /00 0 H e H’ 4 2 ‘“ spon [email protected] lead distribution h sllp .strsarrr 2 feef FIGURE102.-SPRU had dtstrlbutlon InSUP .stmm. Psak loads in powsr-on@-ups E okmg 4 . PRESSURE DISTRIBUTION OF A PW-9 6. In power-on maneuvers involving high angle of attack, a varying inertia load exists aIong the fuselege which is critical for the engine support. This inertia load should be considered in the fuselage analysis conditions. 7. The strip method of computing span-load curves for the high angie of attack condition gives results which check the measured load distribution very satisfactory for wings tapering similarly to those of the PI?’-9. S. The position of the center of pressure at sectiorw near the tip is an important factor requiring further study. In the present case on the upper wing the center of pressure locus bends to the rear at the tip, thus increasing the front spar bending moment in the bay for the high m@e of attack condition by an amount which is not provided for by the present rules concerning tip 10ss. 9. Design rules should include proper determination of the center of pressure position on biplanes. 10. In barrel rolls the maximum dkymrnetry of load occurs shortly after the peak load, and the total load when this maximum value occurs is not appreciably leas than the peak. 11. l’i’km the maximum dissymmet~ of load occurs in a rolI, the lower wing is not shdled and the asymxnetricaI load@ on this wing is of such a nature as to oppose the rolling of the airphuie. 12. In a spin the imide wings are stalled approximately to the plane of symmetry, while the outside wings remain unstalled, aIthough at a high angle of attack. 13. Leading-edge pressures on the wings reach values of the order of 450 pounds per square foot, both in dives and in pulI-ups. 14. Ccmditions governing the critical rear spar loads shouId be studied at greater length. The present report shows that in normally executed pull-outs from fast dives, rear spar loads may be greater than have heretofore been considered. 15. Tad-1oad specifications should be revised upward. This is particularly @e of leading-edge loads, which should be at least doubled. 16. The effect of the dip stream is to cause dissymmetry of load on the two root portions of the lomr wing, and to increase the ~oad on the central portion of the upper wing without giwing rise to any dissymmetry thereon. 17. Abrupt changes iD contour of the wing curvature near the leading edge may seriody affect the lift.. . PURSUIT AMPLANB IX FLIGHT BIBLIOGRAPHY Reference 1: Handbcmk of Instructions for Airplane De&gners. Engineering Division, Amny Air Service (1925). Reference 2: GeneraI Spaaiticatione for the Design of Airplan& for the United States Navy-SIM+B. Bureau of Aerunautios, Navy Department. Reference 3: Airworthiness R8quiremente of Air Commerce Aeronautics Btietin No. 7–A. Aeronaut ica Regulations. Branch, Department of Commerce. Reference 4: Norton, F. H.: Pressure Distribution Over the Wings ofan ME-3 Airplane in FIight. N. A. C. A. Teefsnical Repart No. 193 (1924). Referenoe 5: CkowIey, jr., J. W.: Pressure D~tribution Over in Flight. a Whg and Tail Rib of a VE-7 and of a 7?S Airplane N. A. C. A. Teshnical Report No. 257 (1927). Reference 6: Norton, F. H.: N. A. C. A. Recording Air-Speed Meter. N. A. C. A. Teo@ical Note No. 64 (1921). Reference 7: Reid, H. J. E.: A Study of AirpIane Maneuvers with Speekd Reference to Ang&m Velocities. N. A. C. A. I&ok&al Report No. 155 (1922). Reference 8: Norton, F. H.: N. A. C. A. Control Pc&ion Reeorder. N. A. C. A. Technical Note No. 97 (1922]. Reference 9: Norton, F. H., and ‘iYarner, E. P.: AcceIerameter Design. N. A. C. A. TechnicaI Report No. 100 (1921). Reference 10: Coleman, D. G.: N. A. C. A. Flight-path-angle and Air-Speed Recorder. N. A. C. A. Teohnical Note >-o. 233 (192tl) . Reference 11: Rhode, R. V.: The Premwe Distribution O\-er the Horizontal and Vertic~ Tail Surfaces of the FtiC-4 Pursuit Airplane in Vrolent >laneu~ers. N. A. C. A. Technical Repart No. 307 (1!329). Referenoe 12: Howard, H. B.: Some ProbIems in Aeroplane %ructural Design. The Jaurnal of the Rayal Aeronautical %ciety, ApriI, 1926. of Strength Calculations, Air 31inReferenoe 13: Handbook iatry. Air Publication 970 (His Majesty’s Stationery Office), 192& Reference 14: Laeser, jr., O. E.: Pressure IMributian Tests on PW-9 Wing 3fadels from —lSO thraugh 90° Angle of Attack. N. A. C. A. Technical Report No. 296 (1929). TABLE L-CHARACTERISTICS OF PW-9 ‘~pm:;$~m::&&;~T&zti-H_--: Cantrfc Chm’d of lower W&-—-– DfskimM llmn cedar Ik to eenfrfc_&xd---—.— :%?-_-——— Dihedral UQW W@ lower surface) --= ~ih&mf ~n=WhE Iowermrface).——————— - d’%l& adgeofmot sectfo&loWerWfnG Abre fowar surface* wtfoq Iowerwfng. Distance from C. G. to centerMe of eyevntm --------—Dfstanca horn C. Ct. to cmti Hoe of rudder —— ------------–— &ea of ripperwfnL-__.-–—.———_— &aaof Iow&wfna-.. Totaf wfu area .---_AreaofhdrontnI Mmrfars_ ———----— k of WMC91Ml SnllkxS—–.–––. Wfght of akpkmednrfngte&._.._._R8ted hmapmrer at Znxl r. p. m-.--––– hlEMORIW NATIONAL AERONAUTICAL ADVISORY COMMITTEE LABORATORY, FOR AERO- NAUTICS, LANGLEY 4163c-31~ FIELD, VA., Februa~ 3, 1950. l* . . ---- — — — -.— — -= ~.— -— - .— ——-— : .— .—.— .— - —— -.— .— .— .=.— — :— .-. — AIRPLA.XE spanc4upparwfng_.– Posm lcadfn&---.Wing Ioadfng-. . . ..-––.– LANGLEY Til .—.— .— .— ,— .— ..- .— .— -..— — - ?.92EQUUdaF hp. m Undsparwlmre fwt. M&&+e@ ~ Si5 (approdmafdy). 2#.M (.9ppmSfma@Iy). - — —~ -___ = —.. = ==?== .— .~ 772 REPORT TABLE IL-TRUE NATIONAL OR;I~~ATES Rib OF ADVISORY PRESSURE T tier -— 11.—TRUE ORDINATES RIBS—Continued :tiitlon Per&W# 4.02 :: &6a .W i% -: % –.$ 1;n -. IL 12 ‘a _ I la 62 9.m am 6.28 ;: ’44’ccl Zfif L 14 .E1 ch&64 b!wtippw ~L&+ — Rib upp@r Lawer ;tstfon —. L26 .67 . al #u o :: .Ot o 0 0 0 ; o 0 I !o : & . 0 I Chori=Ei3.S fnchss JPLW ,1Lower .. rmmr .—. —lPer tin :4J w : 0 i o 0. 269 L15 . :fl .lo 0 –.L2 –.14 -.02 0 : 0 4.64 224 .:.81 l.. -.42 —. M 0 0 0 0 -. TABLE [L+veI Olght pra.wms are given 111.—RECORDED Le’ial flight at 74 m. p. h. q~14 ~unds ‘: g. g ‘: ‘: ‘g ~ per squm -------------- 15 ------- ..:_.. -----..Z :1:::: 19 ------- ...-._ ------- . ..-.. - -------------- $ ~ ;! ibi Run h-o. W. ? “i fwt] per SJoarefoot 74 .-..--- ------; i : ~i --–... z::: ..-_. .: . . . . ;::;::: ----- ii z::: ::::::: ::::::: . . . . . . . . . . . . . . -_ . . . . . - i s - Level LUghtat 112m.p. h. Q=”8NPounfs per square fmt I L-. –86 -82 82 41 !...... 44 . -----82 --..--18 ------- -81 2 .--*81 u % i -----.— ~.-.-.-------------.- . . . . -.--_. . .. . . . .--... . . . ---- . . . . . ..[__ . . ------- .-.-... ..-_-. . . ..-_ ..--... . . . . . . -------------. . . . . . . -----. ..-.. - .. —-. ::::: ::::::.:1:: L ,rl -6 -L, !! 89 81 7U 67 a2 18 –ij ii 40 25 ;: 11 ------- . .... ::::Z E __.!!.1:::: ::::;1 ;::1:: :::::;: :::::: E::::: -:1::: ::::::::::: .:::2 ---- —-- -------- ------- ------- . . . ..-. .. . . . . . ------- ------ . . . . . . . . . . . . . . ------- .-. .— +$ –20 —. 11.1 -A -~1“J: :;; -81 —. —.. —. -d -?91 24 .— - -..._. ------- ‘_. E E 18 .j: j -~ i !21 HI ;2 28,-----:-.3. 1“ al . . ----- -----56 i 81 ------ –-..-. @ 40 u .- . . ..- ---. . . . 44 11 . . . . . . . . . . . . . . . . . . . . . 14 -------------------_._.. ------- ------- . . . . . . . . . . . . . . . . . . . Q !! w II — L6 -----------------------------------------...“-.....- - ‘-”-““”” II Run h’o. 28. I.erel fright at 1S.2 m. p. h. q=45.4 pounds p?r sr&re foot E IIa — PRESSURES in per cent f except fn SIISI stream runs; all othem h pxmds Run No. l& $ .&-< unf,Pr.rcm i%~ 426, 7.2s, a2a 9.221 la41” lL@2~ IL21 , la 56 4.6a; ::. :: .a9 o .— +, .269 : 0 0 .67 —.66 Rib I Station m %76 . ml MJ .16 .CS o -. 06 0 11.24 a o o 0 o 0 0 PRESSURE Uppel Uppa !Lowx — —1-. Uppsr Lawel ‘e;r.et? : OF R!b Uppal fmwa Cpp3r FOR AERONAUTICS TABLE cd&4 c&$l”64 Station Gppa COMMTTTEE -----. . .. .. . . . .. . . . .. . . . . ----..... . I ....... ....... .. . . . ....... .—— .._-. . -----....... . . . . ... . -----------.....- ----------.__:. ------- ..--... . . . . .._ . . . . . . . ..- . . ..------- . . . . . . . . --------------- .:::::: ::::::: :::::::::::::: II 1.111 -1 -1::::::::1 - PRESSURE DISTRIBUTION TABLE OF A PW+ PURSUIT 111.-RECORDED AIRPLANE 773 IN FLIGHT PRESSURES-Continued .- Run So. S3. Level M@ at 151.6m. II. h. u=5Q~d9 w I t Tarticd tall SdWe5 Stabilizer and devator Rfb 1 ‘~ iil-ii+il-+w AIB —14i -147 –m –67 66 # S9 4 46 : ?4 9 ------- -91 ii al 19 u ~---l-----~--–--j--------i----- & . Rrm h-o. 6S. Time 1 la ~ al : 3 ~g -c 1 ~ gl %22 9 a : # 31: —i — ~ ; 6 - .“-i I :: ——- 1 -= .- — ,== —----= 0.25aemnd g H -: I _: _: ~~ 16 : _._!. -------- ------~~ ~ ‘~ -----—- . -_-_...—— --——- ----. ---—--— -—--— -—— 8 3 I ; 8 191 a as : ~ PnlI-ap at i9 m. p. h. Time: Owcond Em Xo. 65. I ..— E CD — — — .— square foot 1 --- –;-— — —--— ___: ---: .-. .—--- -- -.- —- ——--- —-—--— -----------—------------- —---–-— —-— I E 1 . — .= ; ‘; ~; ~[ ~- - ; ; [; ‘;”; ‘-~ 1- -.----- .–---.-:------.-_-_.-- ‘; : --—-- ; ‘; ------- --—-- -; ------- ---— ; _-,;; —-- --—– ------: ‘-—- . .—= .- .— -— — L Eon h-o. M. 4i H 16 16 8 ‘HIM 0.i6 serond 47! m 4s 3s 34 ;: as, 22: Is 4 ;_:u: 4 ,..--–--l---—- —-----—..------— +-Il. — 1 –7 :— ,-.-,-- Run Ko. 65. Time: 1 srcond .; 1 ~ c x 0 al : 8 4 6 : 1: 36; 47 47’ 27; 16. g.! !! z z% E 3 53 g 63 M 96 u 20 Q : 4 .—--- g 5 g 80 ls 6 a . 6 22 u 1,------------- “62 66 4: 49 2i M 76 60 37 b2s ----- 1 ~: *I ml 36: RI ~ 1 ‘; Is 4 3 6s 63 39 17 : # ij6! g $443 al ; Isl 23! H ! 24 16 ! 23 3sq11. -. _-_-! -----l------L–- ----—-------------—----- --–-–-.–-–______ -—- ----– --— ---------------‘! i 1 ------7 Run h-o. 6J& ‘Him: ‘1.12seconds : 8 4 ~ — — —- ..= ‘. “+=;+ ---1+k+‘-~- 62 36 % L5 z *r _m— - ,---1-:= :::::=: _ ._—.-q—y -—-.=E= ---.----, –9 -= .— 1 !-..-/–- 11 2 L–-+ t I 1;1 251 — ..— ,= —.— m a -.— .. — H 8 ‘1’ L._= I Run No. .S6. Time: 1.26wconds — 63 46 sl ls 7 ~ ---- --1 — — 42 63 66 a 17 E 16 ---- -— 10 z ------ ---- 26 Is 10 Q ,.----- -. 11 7 s a -: .—— .= 774 REPORT NATIONAL TABLE ADVISORY COMMITTEE 111.-RECORDED FOR AERONAUTICS .-. PRESSURW&Continucd - No. 85. The: Run Uppez wing” —1 R4b l.-ncb ‘Laww Wfng i [ I *~~flbmadelevat~ Ve.rtlmltall uorfam , —. c“ AB I I ... ...- — . U6 Uo 440 40.0 20.0 14.0 0.0 02 67 47 34 1; Oi ~ 2 ~ 67 47 34 16 6 a 8 .65 53 $3 16 4 2 . . . . .. a “49 4i 23 M 7 ; ~i u M i -.....:- i ---.— i -A I : ._._ Hi u . . . . . . . --... -.,—.-. ------- . . .._. ..l -1 ..------- ------:;:- —~1-6 I -9 -1 x ----------..:z::z ----- ._.!. . : ;’6,2 ?I -1 ~1/ _.:! _; -__. . .:-”--------. ------. ___... .-_-.. .._. _ —--- .—-— -— -. ..-: . . . . . . .-— —.- ..-. .------- __ ------- _-::: 1- - ~; - - T Ron No. 130. Pa&up at 11SnL p. h. Tfme: -0.02 second 7- 1 1.1 lti.z 16.1 4.7 ‘“l ‘iZ1 *O . . ------ I I Lo I t; I I I I I 4 .-... - -2.6 E.2 &6 0 ------- -13.6 u.u Zz.u 2L5 ------lU.4 ------5 0.2 %8 2a.4 :: It $ --..--- -----;; -------15.1 ------- .-- —2.6 11 42 4.2 --? -.-.--- .--..-1 –.2 1 .-.:.. ---- ..-. — _-.-& ----. ------- .. —-. -. . . . . . 1- 1..1- –:: ....... ....... ...... ....... -._-. . . . . . . . . . . . -...-.. -..-.. ----. ------ -----------6. i’ -...: . . . . . . . . ------..— .-— ::::: __ –l::::::::::::: ------- . . . . . . . . . . . ..- . . . . . . . . ----- Run No. 1S9, Time:O= .wcond : 8 4 5 i ! 22..0 g: m. 6 g: la o –la 2 10.9 ;4 g“ ti4 g, k: 24.4 Ml %; 17.2 14.0 %: 16.6 c!.Q :: 4.2 .-._ 8.2 0 0 2.! g: H ii’ k: LO ‘2).8 2L8 81.7 :; -.V- _.?-: .--:!. — . 17.7 102. 16.e 140 10.4 ‘“ 2 ‘--i~-~ ------23.4 16.6i% -----6.7 ------- -----—--– ._ _....i 12$! ------- –h”% -.... - –24. 9 ------- –2A 4 –.@. 2 lh 6 –23. c1 =ZL 9 -7. a -H. 4 . . ---- ------- 15.8 [-1L4 -12. o -10.4 -192 I_-=-,. -G ?- I -!. ! -24.4 -&a –u 4 ..-.-.. . . . . . . . ------- .-:-_--... .- . . ..- ____ . . . . ..- .- . . ..- ------- —---- . . . . . ------ . . . . . . . -.._ l-------_ ~_-. -l.. -.. — .------._._ ........ .. ___ ___ . . .._- _.._ .—-.. . ------- --... ~.._~_._.l . . . . . . . ------ . . . . . . . _._.. .- . . ..- ..--... ------- ____ .-_--- .------ ------- ------ . ..-.._ ----.-.------- Run No. 180. Tfme: 0.43second I .1 me Z4. v 15.U 14.6 6.2 I ““o. Ill a... . NJ.4 U. b 6.2 .— ..3 . 1L44 8.8 8.6 7.a I I ~filmml 14.6. . ..1 u. u . . ..1”-”.1” 52.0 8L 2 17.7 a 8 . ----------- #i Og 7 1 -10.4 218 ~&o’ –lo. 9 !E:E=E= “=-1-+ L 7 240 _.... ___. . . . . . a8 -4.2 . –14. o ‘--- ‘“---------—-- --—----------------8L2 -MO ----–14.8 -26 ---::= ~:::: -sao ------------------._-.. ------------.1111 ‘------------------ . . . . . . . _.< 0.76eecond — 1140 10I.o 87.9 H 2L8 6.7 * 109.0 1140 119.0 86.3 lM. o 9Z0 76.4 09.2 67.6 7L 3 82.a 32.8 %8 13.2 E: 2: 16.6 L---6.2 m’o;%g T !! i --.– -1 77.6 Oa7 55.6 82.2 140 l&~ “ 2 “I”ml L-,-------I-3+1 Run No. 130. Time: us.o H 89.8 I 106.0 17a”o 116.b Ill& o 14Ao lm o 119.0 111:0 %7 122.o 117.O 36.8 ~: 67.2 624 &8 40.6 24.4 -.?-:. 9.4 ------- -----%4 II I 724.0 lg: ltiOl 4ao~ 19.2, 15.6, 145,0 12#g la-o W& 42.7 m ti9 13.e 16.1 .-— ~~ 6U8 W. 5 a4 87.& 49.4 43..3 77.0 7.2.6 81.a 80.9 H 86.6 -----19.2 23.4 87.4 41.1 26.0 ------85.9 ,E-i23.4 8.8 ----32.2 . 7.8 . . ----- -----------——— —---- - -------- -- .-A 1 _ 6s.0 —-.– 1 . ..-.. ------- –18. 2 –17. 2 –32 2 -M o -2L 3 47 -9.4 --_-.- ------- ‘---”-------- ‘----------+ ::::1 ----- ::Z ----’--1-. 176.0 125.0 89.0 87.Q 19.2 16.o la 9 Run No. KM. Tfmw L23 wmnde lso. o 13$: ti6 2L 8 19.8 la e 136.0 lx. o 728 42.2 E; 14.6 97.7 IW. o 89.0 o&4 08.4 &5.2 E: 03.0 42.2 68.0 59.8 n7 ------29.0 33. ~ -----JKI.6 69.2 _;;&: M.6 ------4L1 9.9 ..--_– -.-... ------ ---..--- -------- . ..- .._ _.--. . -----------..--... -1L4 -9.3 -L O ------- 69.2 17.7 -8. s -~ 4 -16. I -?. 8 ___ 6s.5 10.8 –6. 2 –~ s –I& 5 -6.8 __. . 44.7 3.1 ------–H. 4 –9 a . . .._. --.._- .. —.. ..--.. .-_. .-_. ------____ ----- -----.- -._... -----. . . . . . . . . .. . . . -----_ .-_.-. . . . . . . . . . . . . . . .– . . . . . . ----- -.... - . . . . . . . -...-. _.-__-... . . . . . . . . . . ..- -----~~~~ ,1 . .. —. . . ..-_ .._ ..-.. . ----____ ___ .- PRESS~ OF A PTV+ DIWEUBUTXON TABLE 111.—RECORDED r PUBSUIT Rfb --l — A 775 .: 1 — ! , , , I ,VVY-, ,“-r-r 65.0 69.s 1 :: “62 040 17.2 U2.o Iaa -a; ---- ; _.2 ---J –r. 3 –AI -31 -1 -----I-J —– —k= -— -- I .- .—— ,= = ..— ..— --— ~.+++ -—- — ..= -—. ‘l-FH=:FI= -2-:;=– -–– –-– ----- -— 520 H H 14.6 = -— _ Rnn No. EM. Time: 1.73ermnde ?aO 63.8 ?3.s 4L 6 19.2 MO &8 — — Stabllh.erand elmatar 69.2 9SL6 5L6 925 M 46.8 421 86.8 ~. 8 67.a ~: .--5.i- _2ao 33.8 ao 49.; %4 33.a tis --2Z4 6t6 –-FE =6 140 -----—El !zl Q 33.3 . fi6 l–=-~—–l--–-–l-––-l--–--l––-.-l---:-!- z: 16.6 9.8 : .- I 1.48aeeonde B 35.8 4L6 4LI ,. IN FLIGHT PRESSURES-Continued Rnn No. 130. TM upper Wfng AIRPLANE := -. ..-.= ~ — “1 80.0 640 6Z: 48.3 g; 22.3 IA2 Ia2 Ii 2 222 – 2L3 –H –t 8 –7. a –2. 6 Z8 I .— .—. –2a 4 -24.4 -63.8 -&6 –K 2 –X 6 Ia. o &8 26.0 l&: as 250 %6 17.2 la41 H 48 ,-_E! ,1 3L2 ! P’ l&6 36.4’ 3L9 ------2aa: ??; 13.2’ 112 .= — ..-= + .Run No. 132. Tlmrc a23 second 1 I 1 -1- ‘. l-----h –1a4 –6.2 –&s –27.6 -aa –L6 -7.8 –5. 2 -–-140 –la. o –la4 –_l ; –~; ; –-6 —---------- . I Run No. 122. ,, Timx 0.48a4c0nd , . -– .<_ .--— . – :— — . - —----- —--— — --— — —— ,, I 1 I ,—— _j__+_f__ = — ~ .- -+- I Rnn No. L32. ‘rime Ma aeumd I 62122 II E: 22 8 u z o 46 .-–; 46 -—-— 1 Run No. H — .- -- ---— Ia - , .— 1 46 47 64 a9 I_ — -——-. —------.—- .— ......— ----.---—[_It ‘rimw aea aewnd I , .- m lW 112 40 _ .-.= -— .Bon No. 122. Tfmu L23 aeeonds 122 a 128 M 59 76 44 —— f! 1 Ma 140 62 30 —-— U8 130 E M t ..— =_ -— .— 776 REPORT NATIONAL TABLE ADVISORY FOR AERONAUTICS COl@l!Fl?EEl 111.-RECORDED —. PRESW’RWContinued — f— -—.. —., . ..- t 1 Rib 8 g I ___l ----___ ----- !::::5:: ~ 4i~ :! ---;f:: ............ “29 . ._-.. 444!2 L------ –9 - -.-:-. ,- . ..-. ..—. — . .. ---- ----. ----- - . . . . L-.. -.-f._... -. . ..---- +:: -. —--- . . . . . . . —— .---.– ---- -- . . ..- . . ..— .-... . . ---- Run”No. 127. Pull-up at 181m. p. h. ‘TImw Oaenond Ron No. 137. I TlmE 0.13seaond t . . -1 Run h’o. 187. Time: 0.38second –06 -8 al 4’a _..?!36 ; 1; –.–- 76 76 .94 –; –a .__::. u ~ ~...-..~.---.-,-._-..-,-...-...,-.--.-,------ ------ -.-... -.-— . . . . ..- 24 .......-.....---—.----.-l-._... –u .-—-- . . . . . . . . . . . . . . -..— ..-.. - -z -34 -.... ~i ~{ ----------- . . . . . . . . . . . . . . . . . . . . --------: -+-: +--: ~ —----- ------- ------ . . . . . . . . ----- .. -. -... :::::::: ..... .. I Rim No. 187, Tfme: 0.76second 2JOI Hsl 841 43, II I I 2441 1;: ..--ti- I ;.’ ~~ l..&_z-l ~ 143I-...x. 161 E4 I I ii-i+ ‘---------------------- Run No. 187. Time: 0.92second 874.0” 206.0 249.0 t ;~g 74.4 403 I a5a W4 266 14Y -. ,----- ‘- ‘-- --”--- I 4!23 39Z0 364 3840 2M Xg.g m 91 ---.2 ----oz. o -.—---. I ‘-–l------l--------l , 1, ,,, ~F=F= -.— -..= .-. -———-. B l--”-------—-------------—-~ ,s l------k------k----L------.L----l------L------L-----l----l---[-------l---- I—-A---A Ron No, 137. Time: 1semnd . 6561 10J, t “-’&l .--—I 817 t 211 m 162 171 . . .----!! n 147 12 ___:Eel ------- i% . . ..--..-—---- II .— ___ . . . .- .,---------------- .,-!H: . . ..:! !! -!- :!;*= =[= E- . * PRESSUFIE DISTRIBUTIOK TABLE OF A PW+J HT.—RECORDED PURSUIT AIRPLANE IN FI.JGHT ..{{[ PRESSURXS-Cc@irIued —— . ===%= --.- Rm Xo. 222. F@t ,,. , 1, barrel rcdlat 167m. p. h. Time: -M5 ,1 k a 1, I .— .— =--...— -_— -— . .— .— — \ second ,,, ! II .< i Is Run Xo. 222. TIrae: Osecond ! _— .- . .— ..— ___ ~— .=..--=--. . .= -— .- ,,, Rnn ‘-- -57 –n ;-_ eLli20za ; =% ~j—---a7; 44 aal 22 U -- .-.., &-O. ZZ Time: 0255econd –91 –!26 17 8 Run Xa 222. Tin.rc 0.60wmnd ,,,1 h ,, ,, , , t,, ,, — ,1,!, -,:-+ -— -= ,— -..__ ..— ~ .=A .— --— .- -. -~.— ..,~ Run A-o. 222. Tin.w 0.75second ..”. — - RUO X-O.222. T!me ,,, 1 1 ,1 ,1 0.SSsecrnrd , 1, .— b 1,, ,, I ,, .“+— — — I 341 29; ~ l-----l-_ .--... -~__------~-----_’h-—--~_-_l Run NO. 222. Th.w: I –6 I semmd –7 I -6 I-_;_ I n! u; ~: ----: --- —-— ..= . -. 77s REPORT NATIONAL TABLE ADVISORY COMMITTEE IIL—RECORDED FOR AERONAUTICS PRE~RES-Continued -. Rnn h’o. m. Upw I I A BIG .- .. . . Tfmo: 1.60ammda — .- —. Rib ,, , -. I Wing IDIE F “;80 .54 42 ,—. . 29 Run No. 222. Tbnw 2 sccda 1111 .:. ..r 1 1 , I I I b“ f 1 I I —”1 l“, 1 I -- Run No. 222. Time: 2.KI mcmds Ron No. 225. Left be.mefroll at 103m. P. h. Time: Osemnd Run No. 223. TlmeJ 0.26swond ,, , # -76 41 29 12 17 [------j Ill 1 J: 4 5: ~ I 7 ‘~ (.--.-::. Ig --------64 ._:. M M’ 36 . . . .. . 16 ! –1: 2s ~---;- 73 36 46 27 la i ------ 72 ------ , 1 .. ,If U ..=iz. 8 ---------24 .-..– –40 -24 -22 -....-. al –19 . 11 -:.?.-::. {] :; -----%7 -34 –lo . ..-. _ ,-_,_ –71 . ..12. —_=- -17 -la . ..-.-. . ..-.-. :-: . . . . . . .t -7 --i -17 -u ‘1’ -9 -8 -6 -.— --; -7 1 -! –8 ........ .— . . . . . ..-.-.. -------1-----:t------l------:t----–l---–l-----11::’:”“::” “E3-=--3’17------ ._:.;:___: 1 :...-.!. 1 “1 ~ 16 ----- ~;:_:.. - .g –s1 -67 -!?8 –la “30 44 48 .,------- - I “4 - 3M : 8 4 6 -62 -g ~ -2d 15 -.-----12i 24 –E9 + I .- 814 .2#fj ‘: -- — mJ 234 182 ~“ E x 331 17 ------19 I ~:~: ::~:::: ----------------- 14 –-—--.-... -----:.. - -: -U .. ‘ ------------------------i---l------l--+----"-l-----II Run No. 226. Time\ 0.76s%ond 119 L1.[ - : mm 211 169 146 .._:w. 144 .J?.% 78 ------83 -----49 % 138 --. . . . ----.-– 17 40 _.2!86 ---— -----..–-.. 20 26 ------- -----------—,E . Run No. 226. Tfnw — –22 –6 “lio -.--...–...;. 81 . ...6. –21 -51 ------–37 –19 –2 –la — —.—. . . .—a. ..-—. .-— &.- 76 ------- .-._ -—.‘ . -. -,-- ---–IS ---_; -9 i –z --.:. -32 -----–la :..... –34 -32 -=% -82 –8 ------–-94 –30 –id .-...-. . ----6 -16 ------ —--- . . . . . . . --. .— .—— . . ----- OM .mxind l!k-.,. 1$ -s .-..-., ------. . ----- a97 ------191 -..fi. 240 llxl -— ..$ Ml : . ..--.. .% aal ma lQ3 .: .al abl 320 137 ;: Wi{ g ~~ I e8~ 136 ------g ._=. & 24! 12 g 242 ~ 161 --?m ,----@ .—-. ;:l:: __:- --Ig + —--- .---:_ --.:: –E .--!:2 -17 ----:g -3 .. —...— ..-. — / _____ -34 ..--; :% _.;;. 41 . --:~- :: .1:-, -17 ‘H –m -------a .--- —--% -13 . . . . . . . 36 d‘“ __!!- 1% :::::l_–- ....-.12z---- ..-. 22 28..--... --... —/––.— -------=: -–Z?...=!-__I!l!. :I;-...l..t-...l .....-t-—t--..--+---4 ,.M ..+ .-. . ; a 4 6 - .— PRESSURE DISTRIBUTION TABLE OF A PW+ PUBSUIT 111.—RECORDED AIRPLANE 779 IN I?LLQH’J! _ PRE9SUREWontinued _— .— . . .= --— — -— . ___ Run No. 22.S. Thn.% L50 e.econds I —. :— ___ .-e - -— ~ ___ - I — — Bun No. 225. TIIIIIX 2 sernnda I I I I II I - .= -L.. “— — ._.. = _._Ron NO. 225. ,,f # ,, ## TiIRe: 8 swonds , ,, .11 1, ,,1, I --. >- Illllllt I 1!1111’1 Eun No.227. h#t — —. — ~ Illli S@L TimIKOsecwd — -. [ .= :..— --— .— --.— I Run 1 h i r, I No.227. TfrJw t I I$wond t m,, –a l— Run h-o. 22?. l%nw la I swonds I %8 ~. %’ : a : : 62.4 47.9 H’ 17.2 IL O &f! m.o 7112 ~; Xt8 10.4 42 67.6 ma 47.8 33.8 l&2 :; 646 a20 32.2 M 6 ;; . —----- 3.6 9.1 620 X4 $; 52 i: 0 -3= .- - - ~ !, 1 El –6 -a -4 -s -s I .- as all ~; 23.0 ‘ I&6 Ih; .— -- 1=4 -6 -s –: — -4-4 –s -s _ –2 —2 .–— Run No. 227. Tim&: L?&saconds t— ------- –9 –a -- —–4 –s —% _: ! –6 I 4 2Z6 2&0 la: j H-: 7a: —— j –; –6 –2 – .6 ,., I H I&o &o –2 ---.. —-- ---- I :A ~:.=i- --l 12 --,- .–- ..—- —.._ —— &..— -~ —.. = I I 12 16 al 65 “:lg u la 12 21 I 19 .__E 8 ~! M .--?: –; –: 6.8 &: ;---= __=__ 9 ------–2 as .-- ——— —-9 0 42,------— ___ ----.-----.- .------ ---— {-l T I - —..= — ..— ..- .— .= -. 780 REPORT NATIONAL — cohfmmlm ADVL50RT FOR ‘AERONAUTICS — TABLE 111.—RECORDED PRE~RES—Continued Rfb _ >.. Run No, 227.” Time: 251 seconde : j : a 69 61 58 : : : ; i; !! ,: 9 7 6 .8 4 _._.!. .:$0 1; 4 ~~ z 25 18 M 21 21 ‘a :;; % 16 --.11 : 16 ----21 ---1: . . . . . . . . . . . . . ..12 ---1; 6 ------- .-.-. ._._2 ------ . ------- ------ - -.—.-.-+.+. . I 1. ~!””m 1; ‘ al 26 lLa ; 4 8 : a ------- ~1 ;g n 2 --;i. 16 -----t. ! 0 ------ -4: . . . . . . . iI ‘ _r4 .-_.h. ~j -4 -s 9 .-_.l ---- .! ~: -.. — -a -.z;. ~ :: 1 — --— t‘ -6 -3 -4 . ....... ........ --------- . . . . . . .1 — -. .= ,.. .— — H3L 34 -..ii19 -------6 2 : -_!—-. _._— Run No. 227. Tlxrm 8.76wconda 85. 28 17 ~. 44 % -------: +!- T .,11”1 17 : 4 81 a6 25 16 9 , ,1: ..-: . ..- -! -9 -6 17 ..- .-. .— —--I -; . ..l.. -8 -6 -4 :: — ---- :1: ....... . -. ‘--- . &—=— --_ ._,-- “--:----:F’= ‘:’=--”*---2--:7===-— a .. Run ~0. 227. Tlmw 4 seconds ‘x 89 –: -8 -4 3!EII T3- ‘--:’ - 34 ; 60-!3 2$ 16 12 : 48:” 39, 21 27 _-?iJ:_.:: s 23 8 b 2 a 9 “5 __ I.-_. ____~ -----------------------—. -— ..---.. L . - ..-— PRESSURE DISTRIBUTION TABLE OF A PW+ IIL-RECORDED AIBHANE PRRSUIT 781 IN PLIGHT ..— — PRESEXJREWonttiued — _ = — = .—..+ .= .-x _— .— .= —— .~ R~ h-o. 227. Time: 4.75seconds I _-, ., —- Ron h-a 22i. Time: 5 ssmnds. . Ei61 El !zi[ Ilkl[ltflllll 17[10]41 ml “51 ‘ 1-------1 61 4 I 71 la I_-_-l la I-–-–J-----I a l------l----l------: Ron 1 . ., RIEhi s@n. ,, r t -- — .— == .— ----- 1!011[9;1 --------l-------l-----i-+---i-;-i----- N-O. 230. , ‘Immm .- TIm& Osecond% . 1- –z –---w’’=l=l---~l--I Rnn No.2W. Tti w 31 22 : 2 28: 3s 20 29 ls 11 2 12 3 5 1 —.. I 39 al 23 10 4 4 1 la 1 5 16 ; E –--–i-: ---7; E? 2, 10 -------5 ! 2 –--–. -—– 4 –1 -1 . ; -- ---- .-. —-- ,-----—- —-— ~.-—..A I al $ 19 13 10 ;: g -1-+ Isexd. I 16 m 19 16 la 6 4 --:= -. —— ..— .L —- 9 %8 Run No. 230. Tb.uc l.ia wends s a H L 40 1 9 % ‘–-_%4 _: ; : 4 5 0 -— -------—--- i –: ::–6 –3.o –3 –. 5 -a :. ------ -—- 80 5— —- 2_ 0 — _= - — I Run h-a 2iKL TIIW 2 srmnds. 1111 la k: I –: -----–4 –2 –3 –. 5 6 .--=i- s i –! a4 o :5 II — .-_.[–--,—-[.---–[-----, : 25 Ia u 7 0 a3 –: a 8 4 2 -L 3 .;0 : 1 ‘-- — T1 3 3 4 a 3 6 -_-! 4 --1 -—.— -—— —— .-— 782 REPORT NATIONAL TABLE DvlsoRY CObiMITITlli 111.—RECORDED FOR —. .__ AERONA~ICS .- PRESSURE$-~ontinued ,=_. .. .,-- Rnn No. 230. Time: 8 seconds, . I I f I I II 1“1”1 1 1 “-1111 I ,11 t ‘ I ..— Run No. ~. I TImw 8.76s-wends Run No. 230. Tim% 4 wndF 1 {11... r @ 42 m u g “11. 1 21 24 a H 10 ti 4 ------- 16 17 10 12 I I ~ 23 11 8 i! IO 1 r t 32 u z 19 . . ...?. i 18 9 ------ ,11” 12 1 ~ -: –4 --- –4 -10 ..1-?. 40 2 4 I 1 I .! i! 4 -..--6. ; h 0 -3 0 ! ; 7 .—:?-. f ;::--- ----”-- ‘---- -[-------1----1-+-+: 8 l------l-----:--l--------l-------l-- I s ----4 -0 8 ~. Q “y :---- :“:1:: Run NtJ.233, Time: 4,6sewde r 39 ; j : s 4 : T “82 83 78 ii 42 !23 12 10 63 % 1: ~ ; 8 26 y 19 7 ..-..-:- 23 20 16 ! 22 21 36 80 u 14 ii ? --: — 62 “] %ti 3 ..–-–. 21 13 ..--! Hi ‘:---i..-.. ‘-----E-— ------ _; .? -4 -----:—— —--: = -.1_ 2 ------- ; --? --: –8 : 18 $“ 9 64 % E a) 25 6 M 7 -.-.: w !23 f: 21 18 10 H 2 18 29 8 ..-:. %. 50 26 61 24 30 65 z 17 29 18 36 -------lQ 17 lb -----13 .--?.-..?“6 ------_.:- ---:- ::::1: 1:? --”Z[ -j $ 8 ; 7 .-:L------- 1 I .— 79 ?4 ~; ‘; — .-- —-. ‘: --::‘-----” .--— .— ---- ---J-- —.. ‘---F -:----- -------; . . . . . f. . --..—. . Rum No. 230. Time: 4.9 S?COdS 10I ‘–5 8 –: –8 -s 1 --:!. 1%! -.6 -M. o –6.0 —2,o ... 40 14 -! -4 -2 –---l---=. .. — Run No. 230. Tfme: 6.25seconds I —.—-. —.. . -. PEESSURE OF A PW+J DISTRD3UTION T.4BLE 111.—RECORDED &tKPL&NE PUESUIT 783 IN I?LIGHT “-: l?R=~%Contiued - Run No. W. t , -Verucd tall Sorhca upperwing ~tTl:lL c!DlE[”F16 — Timu 6M SSCOndS 1“ * — .— — * Run No. 230. fie: g-11:1 ----- g L — 69 !m i% .— 20 la - s 5 : 23 26 Z 16 6 seconds : a la .—; : –---i- a i—-– U -— — —— i . -. I TmIT a_ la ;8 ~a 1; 9 8 a : -lo a : : _; –: “--—____a .—— ----- —- — 6 7 1 4 4 —-— 9 :: --— -.—. ---— -- —-. —. .— .--.-Run 1; 6 i-. 11471 ___l M[ No. 230. TimK I 7 eeconde W-PI==~~ t–-i–--i--—t—t—t— —. .= 10[–.----l–-—-.’--–--–~-l-—-. Ron No. 2U. Inmrted Wht at 79m. p. h -2 -6 I 6 ---:—--—_5 -— ~: — –9 s --3 -6 -4 z -2 ------ .----— Rrin No.226. R 1 8; a z o~ –158 -E ---U4 6 7 --?- Rtm No.2?6. ‘-—-—------ —.—-— ------ --—— : – L — –aaa -W –9 2 –9 .-— ‘----- -aL –40 -7 –7 –-i .— 5 . -..---- _L-_ —— —- --— ----.- .--— . -— —- -–=i- _I- .2----- E I — -— - ------- -—— Polloutfromdke –n.o -18.0 –26 a.a -U EKO No. 216. Pueh down iktMl m. p. h. -m —“ !1T -75 —: 6 -la , ..---la -—— 18’ H::i 7; : –laa –: -198 –lcO -83 10 –a 11 M 12 -_l 25 : —--- -7 ;: 4 1 ..— -— —-“13+ .. ..— -- ------—— 11 --------- —----— .1 8 12 A-———— 1--1-1 -— 1 I 10 —–----– _—---— T 9 ---- h ! -149 -~ -; -3in a 46 Is -----— -11 — ----—— ‘r ..,— .< . i ‘k=+L:-““ -I.I — Dlvec.t~mp. –149 -aia -894’ –14s -la4 --20 -10 -6 17 a2 +%= . _~ -— -- -— ----i= -Zw —e .= . — - —---6— -i .--— ------—-—.— 1 —— ---— -- —- ——.— I_ — -—— I -— :-G -A >— .— -—= ..= =-.-+- REPORT NATIONAL TABLE ADVISORY IV.—RECORDED COMMITTEE FOR PRESSURES [Pwnds.E —. AERONAUTICS ..—— IN SLIPSTREAM square foot] -— Mn No.MM.Lavel .1” 16 1 “. : . 11 ~ 1.. , — 34 21 al 86 44 -.A;_ 28 Ie 11 6 a 26 84 27 1: E 2U 26 17 1: 26 17 — fllght at 82.7 m, 4 p. h. ii T WIXIY ,., -E? –4 o 4 6 1 –6 22 29 19 Ia J Ii ;! 4 o 0 7 –k 0 : LO 8 :1 I I o .. . . . . . . . -------- ------- —.--. ..- —-. — : I -....-: --------- . ... Run No. 332, Level flight at 114.6m. p. h. Run No. 80S. Leveltlight at lbs.7nL p. h. 8 w z 0 I : ! 7 -89 –: 44 70 22 20 -7a % ~~ 40 M M 8 –81 –8 86 6 84. 22 0 –25 –78 –m 2 21 42 86 18 9 .--7 70 . ..-.-%.. 41 ~ 28 11 -81 -88 29 40 fi. -124 –92 -52.0 -48.0 –89 –la -8 -3 7 Ii: 24 2a 9 &o a –2. o 11 . . . . . . . . . ______ --------- -a? –22 8 z 1 . . . . . . ..- –aa –b 9 9 ._._.!. ______ Run No. 814 Power-on pu[Iup at SLb m. p. h. 8 % 0 : a ; 6 7 M 49 WI 109 84 78 1.82 ti m 15 ~ as 15 8 @l aa ; 6 107 Ia4 z :+ 22 19 la # .- 69 —--—81 -–:4;. 80 ~ E % 44 24 .20 –-2 20 –-: -11 17 22 22 b lz .--- .-—-. --------10” —-:--------.1 E ~ T 40 –: –10 a -~; :; .__::. .---. — .-----_- Rnn No. 315. PowW-on pulIuII at 114,5m, p.”h. --..——. — Run No, 222. PuI1-upat 75m. p. h. l’imw Osewmd a 4.0 0 o –: : .— Run No. 822. Timw 1 1 1 1 0 -L-B -40 —— -.. . ------ : 2 8 0 0 0 ;.o .5 ..-_ .... I.aecand no as 2a M a7 al 14 ;; 14 0 8 .---. ..—. .- . .. ---- ..-..—-- ------- I ......... I +--------” .— PRESSURE TABLE I DISTRIBUTION” OF A PW+ K.—RECORDED EM No. W.. PURSUIT PRESSURES PdkP at Ifi5 m. P. h. WIANE 785 ET PLIGHT IX SLIPSTRE-lM-Continued The: Llsecond i .. ..— .~ ...— .—— I I Run ?Ja. 823. . Tim: 1% secmti -- ,— — — .-..— I ~ ..-= — .—— — . ~ _.= — 7 ——_ .—— :~ ..— ..— — .= -- .—-. r— -—.= .— ...— ..~ .- .=–~= Run Nn. 823. Dire at high spwd I t . .-—.. —.— .— .— ..~ = 786 REPORT NATSONAL ADVISORY TABLE COIWNITTEEI V.—FABRIC FOR AERONAUTICS PRESSURES [Pounds pSr SQUSrefOOt] --- I Rap No. W Leveddigh~at 82.7m. p. h, I Orlfim . .— -. — . .. . . . .. I Mvsl flfght at 91.7m. Run No. sol. : -.-+ p. h. I .-----------.-“: ------I-----------i’=”= =a+k::: =+-------d 18 +-L .__.? !-- -+ 1“ -— --- Run No. S02. Level tlfght at,114.6m. p. h. : -lo ‘a 16 .20 14 4 .-. —...------.-— .--.-.-— -------------------8 .-+ -8 6 ..–...-– .......... –d------ . Run No. 803. ; ..-:;.. 26 -.--.-.-.---!-0 .d!.. . . . . . . ..- –——-. I . . . . ,-------------- .–---!. ..–.2 t ,---------- :1 . . . . . . . . Lerel flfght at 184m.p.h. ---% _- ___ .. ::.. =+1------ ‘ 11 !-.-Q. 16 . . . . . . . . . . . . . . . ..1 ,. .,:. -. I h. . .... . ::+i_ ~_.2.. .---. . . . . . :----- .._;; . t ““’~~:),,=~,,1=,,~, Le~eIfUght at149.4mp 14 –88 -6 la “a : ..--. .— -- --. —---- . . . . . . ..- ------- . . . . . . . . . +2 . . . . . . . ..- . ----------—------------–81 I --..+..J ..- 4 14 .- .--—.----------.—-.– g RwINo.804. ; . 7 ..........f ,’” I Run No. 805. LeYal flfght at Iti3.7m. p h. J -— -+ .- i.-.. - Ran No. I Ill’-””ll’ ~’ ..-. 816. Pull-uP, powsr~n, at 149.am. 244 195 1?7 m 6!2 - . . .. —---- -, —-------- .-. . —---------------lea I@ . . . . . . . . . H-------.–—--------- .2a 16 -..-- n 42 .’”’--’”’ ! p.h. .. —--—--. _.. -._...l...-..K. t-l 84 . . . . . . . . . ..--_ I 14 ......... - ..1 . . ..- . . . ..._..?I I Run No. aZ3. DIreathfghspsd I ; :1o7 --------148 –18 –a8 . . . . . . ..- .- —-----–Ea -..-----– 01 25 .---. ..—. —-. —.. -.-..------—-:I -$ -------- .2+ _-%, ,. -.-, I I I:::;6:}::,~._l~_.:::: —-” 6 — -. PRESSURE DISTRIBUTION OF A PIV-9 PURSUIT TABIJE VI.—TAIL AIRPLANE 7s7 IN FLIGHT SUMMARY I I Total Ka2hnum Aver#a ti%i?%f% quara foot ocatIorl { mnx- imum mum I I K31 Pdl -ZM.o 116 up . . . . . . . . . . ...-! ! m ---do— ------------ 132 . --do-— L23 .--do-_ .-.. 1%.: I J ----– ..._do .. ..-... ---_-_l 1ss ,__do . . . _________ 136 ,-...dm------------ 137 –13.8 -41.6 6L4 149 -3i9.o –li?o. o S&o –21.3 -10.4 3.? -43.0 –46. o @lo Ii-4 -243.0 -21L O 162.0 -a 9 –14. I 10.8 -s4.0 –66. o 137.0 N-4 –m –1 4 $8, –eo.o Wo 1163; o Ii7. O 172 I 181 -8.0 3.8 3.2 : , Dire. - . . . . . . . . ..-... -.l-——-- 2M PtMbdovrm.- . . . ...--! N-I I I — .—— — -— ----—-- I L__ —Jl-— 1 ;% 2s o -39s. ;H ~ _m4 1“ -“}f t ,g , -!M.4 -m” –8.4 , -ml I –24 no o i“ -07.0 , 04 N-I -— ---.249.6 -2a: -—— -a. 4 I -=-t-== I - —-—-----— ;––- x-l x-1 V-I $$ x-I ‘L------ — ‘ –::% \ -%.! –12. 8 ; :1387 -40.0 —1— v-1 w-1 x-l x-1 I I R-4 .- .- —-— ! -—~~ 3 _L— .. ___ ‘1 ‘s - ‘i--l -._-.. —. ___ . . --: ..-. -.-.:= .- —.-.--. — -—-------------- --—---- ‘i:_i--*_4-l recorded on Ieft hafsontd W amrfacm. o 41630-31-51 ! 13 -=— .—..— -- : -------E:*{’’:+p~-:l x-l y——,——-—-+----- : P~ -— I 7 --K+-_31#!_’ ‘;O a-l t 149.8 .—— ...— , ---E-; ~; -i!q--E-@?- Ti- -.. .—. -: R-4 111.6 7.5 + I t –43i.o , -33.7 161?fs ~ lM.5 given timing Iiie. —. i-:----”T+l-–- --! ~!--:yf: -1.: —.--—_ —— x-4 ,--.-—---- —_—_k-_— X4 : A& spoafd — --t -is. o M.o .~ 0-1 Pouup. . . . . . . . ...-+ —-----., --- .. . :.= ..-.. ~ -.~ :—. -.— N-2 ‘.-...do..-----. S ,.- PuU-Out fmm dfve.... ~ t 233.0 2316 ------- —- 226 1316 — ‘T-k I Die_ . . . .._--_. 12W. O -— —_.. --------—— A - . - ..-: ..- ...— ... ~:+.z.+-—— — . ..— .,-~ ..:.,, .-:~ _ --~ ,—. - — —.— —-— ?--!..-—— ----.........- ---. —-. r—______ i L — ---y ;—-------. —__— —, ~.... ...--------— !_____ 1 . . . . . . . . ..- m _.. -—-— &l % dl.....--—-—–_i sm ------------------------- —-—. G4 WI —I---—-ti. hit >-— N+ :; 22s ----- -——— ..-. .— —~ --------- ?s-1 -Lo& o 70.0 ! I 186.1 -——----—--—- “.—-------- N4 N4 N-1 -12.0 213 .. -—----------.-- —-— . --------- 1 -453.0 –256.o iz4.o %9 m H*k’ .-~ —--.— .-. -— -- M mhw hum : Wllnti ; aqlmn?foot PrGssnre %.% k; 1~ “---do----------------l N-1 K% .- I PldI up, Power off..---: N+ nmml -276.0 [ 200 0+ -Xl&o 48.0 164 137 ,. . ..~o . . ------~ –Z&z -g : NY N-1 –13.3 1 134 ?: o-4 –MO. o –ma. o S7.o M ----------- -327 –?3.8 ?%: -IL7 –8. 8 –U4.o I ,--._= — I“i Arva#m mxnant Wxtmum Lacatfon shut ! PKsmlrQ of w- ; Totaf ~Apprcdmate sIrwJE@. —. -. —