MftC(V l-<|3 ARE No. IA.F05 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED June l ^ 1 * as Advance Restricted Report L4F05 FLIGHT STUDIES OF THE HORIZONTAL -TAIL LOADS EXPERIENCED BY A MODERN PURSUIT AIRPLANE IN ABRUPT MANEUVERS By Flight Research Maneuvers Section Langley Memorial Aeronautical Lahoratory Langley Field, Va. NACA \ ■ WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- nically edited. All have been reproduced without change in order to expedite general distribution. 93 DOCUMENTS DEPARTMENT Digitized by the Internet Archive in 2011 with funding from University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation http://www.archive.org/details/flightstudiesofhOOIang NACA ARR No. 1J4FO5 1 ATIONAL ADVISORY COMMITTEE FOR AERONAUTICS ADVANCE RESTRICTEI REPORT flight studies of the horizontal-tail loads experienced by a modern pursuit airplane in abruft maneuvers By Flight Research Maneuvers Section s : &RY Flight measurements were made on a modern mrsuit airplane to determine + pproxirrate magnitude of the horizontal tail loads :n accelerated flight. In these flight measurements, . ressures at a fev re used a? an index of th< t 11 loads by correlating these pres- sures with co: r ^lete pressure-distribution ata obtained in the NACA full-scale tunnel. In addition, strain gages and motion pictures of tail deflections were used to explore the general nature and order of magnitude of the fluctuating tail loads in accelerated stalls. The results indicated that, if the airplane were not stalled, a total up load of 57^ pounds would be experienced on the horizontal tail in an Cg pull-up and that, with power on, this load would be distributed unsymme trie ally with about ids more up load on the left stabilizer than on the right. When stallii occurred there v/ns an initial abrupt increase in the up tail load of the order of 1 rcent of the previous load, which was followed by repeated load and stress variations due to tail buffeting. Under the conditions of tail buf , the possibility of excessive stresses due to resonance was indicated. INTRODUCTION As a result of numerous tail failures of modern h ." - 3d ?irpianes in flight, a flight investigation was undertaken tc determine the general nature of hori- tail loads ex rienced in abrupt pull-up maneuvers NACA ARR No. LljF05 Tests were made by the MCA at Langley Field, Va. during the spring end summer of 19h2. The flight-test procedure involved the use of pressure measurements made at a few points on the horizontal tail, which were correlated with complete pressure-distribution data from the NACA full- scale tunny] to determine the approximate tail loads. This procedure gave satisfactory results except when applied to stalls wherein abnormally high fluctuating pressures, corresponding to tail buffeting, were experienced. In order to help establish the significance of the peal; pressures recorded, a strain gage capable of following the load fluctuations was installed on the stabilizer; motion-picture cameras were installed later to record the deflection of the horizontal-tail surfaces. The results of the tail-load measurements obtained are discussed in two main parts. One part pertains to the more or less steady loads experienced in maneuvers, for which the determination of loads by means of the measured pressures is fairly straightforward. The second part deals with the fluctuating loads experienced in stalled flight wherein the significance of the measured pressures was difficult to establish. For this second case, the main dependence is placed on strain measure- ments and photographs of the tail deflections. DESCRIPTION OF AIRPLANE AND APPARATUS Test airplane .- The tail-load tests were made on a modern pursuit airplane having the plan form and dimen- sions shown in figure 1. The gross weight of the airplane was maintained between 11,000 pounds and 12,000 pounds for the tests. The center-of-gravity position was maintained between 20.8 percent and 30. 2 percent moan aerodynamic chord . ■ : "'3s ic flight instrum ents.- Airspeed, elevator angle, sticl: force, and normal accelerations were recorded during the tests by standard NACA recording instruments. The airspeed recorder wan connected to an NACA swiveling static head located 1 chord length ahead of the right wing tip and to a shielded total head mounted on the airsoeed boom. NACA ARR No. li!.F05 3 Pres sure -distribution j rista i lati^n . - Four pairs of orifices were installed on the horizontal stabilizer to measure the pressure difference between the upper and lower surfaces of the stabilizer. The spanwise and chord- wise locations of the orifices were chosen to correspond with particular orifices used in the pressure-distribution measurements made in the NACA full-scale tunnel. A sketch showing the location of the orifices used in the flight tests is given in figure 2. Pressures were recorded for the individual orifices by an MACA mechanical manometer mounted in the baggage compartment of the air- plane. The inboard orifices were connected to high- frequency pressure recorders to permit a study of the pressure fluctuations at the stall. Tail-deflection apparatus.- The deflections of the horizontal tail under load were measured by photographing the tail with two l6-millimeter motion-picture cameras mounted, one on each side of the fuselage, in the inter- cooler exit ducts. The cameras were synchronized by timing lj hts operated by a master timer that also synchronized all the recording instruments in the air- plane, ""argets '-ere painted on the tail plane to identify the spanwise ] osition in t". Lotographic records. The camera installation and the targets on the horizontal tail are shown by photographs in figures 3( a ) and 3(b), respectively. St rain-rao-c installation .- An electrical strain gage was installed on the skin above the rear spar on the right horizontal stabilizer. A photograph showing the location of the strain gage rid the dummy gage on the horizontal tail is given in figure Ll. The orifices on the upper surface of the tail and the lead from the orifices on the lower surface are also shown in figure ip. For one flight, de Forest scratch-type strain gages were mounted along the front spar on the upper skin of the left stabilizer at 3k, 60, and 7I+.5 inches from the stabilizer tip. The gages were mounted by gluing the gage target and scratch arm to the skin. k NACA ARR No. Ll|P05 TEST PROCEDURE The types of tests and records obtained are summa- rized in the following table: Type of lie cords obtained Plight Basic Pressure Strain iai_i. maneuver flight distri- gage deflection bution ll*3 Abrupt pull-ups Yes Ye s i No No 15B Abrxxpt pull-ups Yes Yes No No 18b Abrn.pt pull-ups Yes Ye s Yes No 193 1S0° turns Ye s Ye s Ye s No 213 Abrupt pull-ups ar.d l80° turn Yes Ye s Ye s Ye s 2)_|.3 Abrupt pull-ups Yes Yes Yes Ye s combination of pressure orifices and a strain gage and, finally, to an installation which simultaneously measured the pressure, strain, and tail deflection. The strain gage was installed to facilitate an interpretation of the pressure fluctuations experienced on the horizontal tail at and beyond maximum lift of the wing in the pull- ups. The apparatus for measuring tail deflection was subsequently added in an effort to obtain additional data on the motion of the tail following the wing stall for correlation with the pressure fluctuations and the strain me a s ur em e nt s . The abrupt pull-ups to maximum lift were made at various speeds, from the minimum speed of the airplane to an indicated airspeed of approximately 2ll;. miles per hour. The corresponding normal accelerations experienced ranged from Ig to l|.5g. All tests were made at an alti- tude of approximately 6000 feet and, except for one power-off run, with the engine operating at 21^50 rpm and 27 inches of mercury manifold pressure. MCA ARR No. l4P05 DETERMINATION OF TAIL LOADS The pressure data recorded in flight were converted to ta.il loads from the pressure-distribution data for the tail plane obtained in the NACA full-scale tunnel. Because of an unsymmetrical flow in the full-scale-tunnel tests, the load on the tail, as indicated by integration of the measured pressures, was unsymmetrical. The dissymmetry of lead is shown in figure 5> which is a plot of the spanwise distribution of load on the hori- zontal tail. The variable c n c us^u In this figure is the product of the section crmnl-force coefficient c n and the local chord c. The normal-force coefficients C-w for each half of the tail were plotted in figure 6 as a function of the pressure coefficient Ap/'i, in which Ap is the dif- ference between the pressures on the upper and lower surfaces of the tail plane at the two spanwise stations where orifices were located in the flight-test Installa- tion and q is the dynamic pressure. Th^ tail loads computed from pressures measured at the individual orifices therefore assume a symmetrical tail load with a load distribution similar to that obtained in the full- scale-tunnel tests. The normal-force coefficients for the tail are noted to be proportional to the pressure difference across the tail plane and are also a function of. the elevator an^le 5 e • The tunnel data for the right inboard orifice were considered too inconsistent for use in evaluating the tail loads (see fig. 6) and the evalua- tion of tail loads for the flight bests was therefore based on measurements at the other three stations. Tail loads were determined from the tail-deflection data by means of the influence line shown in figure 7 and the spanwise load, distribution of figure 5. The influence line was obtained experimentally by applying unit up loads at the indicated spanwise points, whereas the spanwise load distribution was taken from NACA full- scale -tunnel data. The tali load per inch stabilizer deflection is obtained by the summation b/2 ^> yw A> NACA ARR No. L^F05 in which w is the running load at a spanwise point, y is the ordinate of the influence line at the same point,, and b is the span of the horizontal tail. This summation shows a load of 875 pounds per inch tip deflec- tion on the right stabilizer and 97° pounds per inch tip deflection on the left stabilizer. Some question may be raised as to how the spanwise load distribution (fig. 5) should be faired across the fuselage, but consideration of possible changes would not materially alter the loads as measured by tip deflec- tion. RESULTS AKD DISCUSSION Loads in uns tailed flight .- The tail loads in accelerated flight were measured in pull-ups to maximum lift of the wing. Time histories of airspeed, normal acceleration, elevator position, and elevator stick force for three typical pull-ups of varying acceleration are presented in figure 8, The present discussion is limited to the loads attained before the wing stalled, that is, to the portion of the maneuver prior to tail buffeting, as is indicated by the fluctuating normal-acceleration curve . The pressure coefficients Ap/q for the four span- wise points are listed in table I. The corresponding values of normal-force coefficient Cy obtained by reference to figure 6 are also listed for the three stations at which satisfactory calibrations were available. Total tail loads corresponding to the normal-force coef- ficients of table I (tail load equals 55<1 g n) have been plotted in figure 9 as a function of normal acceleration. Extrapolating these data indicates that an up load of about 5700 pounds would be experienced at an acceleration of 8g. In consideration of these tail loads, a study was made to learn the contribution to the load of each of the following factors; (a) Increment of tail load necessary to balance pitching moment of wing -fuse lage - propeller combination No. l!jP05 7 (b ) Increment of tail 3oad due to horizontal location of center of gravity with respect to aerodynamic center of wing-fuselage- pr ope Her combination fc) Increment of tail load due to manipulation of elevator . At the speeds investigated, the increment of bail load due to factor 'a) (a down load) was found to be relatively small, about 5»5 C J or 5'60 pounds at 200 miles per hour. At diving speeds, however, this increment is large enough to be of primary consideration. The increment of tail load due to factor (b) is always r m up load at positive lifts with the conventional wing and tail arrangement; if the aerodynamic center of the wing-fuselage-propeller combination is known, determining this increment of tail load for any center- of -gravity position, gross weight, and normal accelera- tion resolves into a simple moment problem. The increment of call load varies directly as the product of the gross weight and normal acceleration and varies linearly with center-of -gravity location; that is, this increment of tail load will be zero for every flight condition if the center of gravity and aerodynamic center are coincident and will increase as the center of gravity moves rearward. gull-scale-tunnel tests indicate that the aerodynamic center of the fuselage-wing-propeller combination (power on) of the airplane tested is at approximately 15 percent of the mean aerodynamic chord. With this aerodynamic center, the increments of tail load calculated by the method suggested are in substantial agreement with tail loads obtained from flight-test data. The tail loads experienced during acceleration were considerably larger than the loads indicated by standard design practice because the propeller and fuselage caused the aerodynamic center to move farther forward than had been anticipated. A discussion of the effect on the tail loads of factor (c) (elevator manipulation) requires a knowledge of the control movement during the maneuver. It is apparent from figure 8 that the elevator force is relaxed before the maximum acceleration is reached and as a result the stick force is approximately zero at the time of maximum acceleration. When the elevator stick force 8 NACA ARR No. li^P05 is zero, the elevator is floating, and the tail-load. increment due to a combination of factors (b) and (c) is equal to that obtained in a similar maneuver, elevator fixed, with the center of gravity at the point giving zero stick-free stability. Computed on this basis, the up tail load due to releasing the elevator is 130 pounds per g of normal acceleration. Extrapolation of the data in figure 10, which is discussed subsequently, corrobo- rates experimentally this calculated load increment. This load increment is indicated by the difference between the curves shorn for elevator floating and elevator fixed as determined from unstalled pull-ups and steady turns, respectively. Pull-ups to maximum lift and unstalled pull-ups to the same acceleration gave dissimilar tail-loading condi- tions. Analysis of the data indicates that the lead was unequally distributed between the right and left stabi- lizers during unstalled pull-ups, as shown in figure 10. The total tail load, however, was the same as that obtained in pull-ups to maximum lift. (Compare 1+.53 pull-ups in figs. 9 anc ^ 10.) A clue to the probable cause of the asymmetric load is obtained by a study of the time histories of figures. 11 and 12. A turn with power en is shown in figure 11. Immediately before this turn was entered, the load on the left stabilizer was greater than that on the right stabilizer and remained greater by about the same amount throughout the turn. The ores sure changes that occurred during the turn were very similar on both sides of the tail and occurred simultaneously with acceleration changes. For the turn of figure 11, which was executed with rower off, the load.3 were nearly equal on both stabilizers, with the pressure orifices indicating a slightly larger tail load on the right stabilizer. The changes in pressure during this turn were similar to the changes that occurred in the power-on turn. Consideration of the magnitude of the dissymmetry in loading indicates that the unsymmet- ricai tail loading is f'ttributaole to a slipstream twist which ircreases the ancle of attack on the left stabi- lizer 2" or 3 in & positive direction and decreases the angle of attack on the right stabilizer by an equal amount . It appears from these data that the slipstream twist with power on Is responsible for an asymmetric tail-load increment except at maximum lift. (See fig. 9-) T be dissymmetry, which is independent of s'.ced and acceleration, 1IACA ARR No. LJ.1PO5 results in an up load on the left stabilizer 800 pounds greater than that on the right stabilizer. This unsymmetrical loading, if attained in an accelerated pull-up of 8g, would result in a tail load of 325 pounds on the left half of the tail or in a stress due to an equivalent uniform tail load of 65OO pounds. Lo ads duri ng sta lled flight .- In abrupt pull-ups to maximum lift, large and erratic tail-load increments were indicated by sharp pressure rises immediately after the stall occurred. The initial peak pressures were followed by fluctuating pressures throughout the period of stalled flight. Time histories of pull-ups to maximum lift (figs. 13 and !.' N show the nature of these pressure rises and fluctuations, together with simultaneous records of strain as indicated by the electrical strain gage. These abrupt pressure rises and fluctuations are ascribed to fluctuations in direction of rbe air flow at the tail, which are due to stalling of the wing. was previously mentioned, cameras were installed to record the motion of the horizontal tail during pull- ups. The accuracy of measurements of leading-edge deflections on the l6-millimet r film is believed to be within ±0.0005 inch, which is equivalent to ±0.1 inch of actual tail deflection. Although a camera speed of approximately 6k frames per second was used, the frequency of the tail vibrations was such that the maximum ampli- tude of the motion of the tail was not necessarily defined. The data were therefore plotted (figs. 15, 16, and 17) in the form of instantaneous beam-deflection diagrams at time increments of approximately 0.017 second during the stalled part of the pull-up. In these figures, if a line faired through the spanwise points at which deflections were measured did not pass through zero deflection at the center line of the tail (see 2.S00 seconds, fig. 15 ) , the beam diagram was arbitrarily shifted so that the deflec- tion at the center line was zero. The shifted beam curves appear in the figures as dashed lines. This shift of the beam curve is considered justifiable on the basis that vibration in the airplane may have caused slight shifting of the cameras or that the zero reading for the particular frame may have been in error; either of these factors would have caused a uniform shift of the beam line. The change in tail load, which is indicated by the deflection of each stabilizer tip is listed, at the end of each beam curve. In figures l6 and 17, the total load 10 SAGA ARR No. LI4.FO5 change for each beam diagram Is tabulated at the center line. Deflections o^ the stabilizer are also plotted as time histories, together with airspeed, accelerations, pressure, and electrical strain-gage records in figures l8 to 20. A marked twisting action of the fuselage may be noted during the stalled portion of the pull-ups. The deflections of the right- and left- stabilizer tips are not, therefore, a reliable indication of the individual loads developed on the right and left stabilizers except during the first part of the maneuvers before the twisting of the fuselage was set up. The axes for the pressure and electric strain-gage records were so drawn that the ordinate s at the beginning of the run and at the time cf maxim-am acceleration are proportional to the loads computed at these points. Because both the electric strain gage and the pressure capsule have straight -line calibrations, succeeding peaks are also proportional to the tail load. The three de Forest strain gages mounted on the left stabilizer provided a measure of stress on the upper skin of the left stabilizer during the runs of figures l6 and 17. The de Forest strain-gage records are shown in figure 21 and a photomicrograph of a typical record is shewn in figure 22. Mtnough a history of the stress encountered was recorded by a de Forest scratch gage, no time record is available. The peak stresses, therefore, dc not indicate the frequency of the applied load and must be interpreted in conjunction with other records. The change in load from the level-flight condition to the joint of maximum acceleration that occurred immediately before the stall is indicated by ALi in figure 13 and the change in load indicated by the first peak on the pressure or strain-gage record after the stall occurred is indicated by Alg . The ratios of the load immediately after the stall to the load before the stall AL2/AL-1 as indicated by pressure-orifice and electric-strain-gage records, as well as similar ratios determined from the tip-deflection ana de Forest strain- gage records, are listed in the following table; ITACA No. LJ4FO5 11 Load ratio, AL 2 /iL! Pressure orifice Electri- cal strain gage (root of right stabili- zer ) Ti defle 1 p !de Forest strain z t Ion gages, from left tip Figure Right in- board Left in- board Right tip Left tip Ik in. 60 in. 34 in. 14 18 19, 21 20, 21 1.5 1.5 1.1 1.2 1.3 1.9 2.6 1.4 ft i b 1.6 J- • J i7e 1.5 2.0 1.0 1.1 1.0 1.1 1.3 i7e 1.3 m he tabulated data show that immediately after the stall a large and abrupt increase in the up tail load occurred. Although changes in load indicated by each of the records obtained are listed in this table, the indi- cations of the pressure orifices are discounted, not only because of uncertainty regarding the dynamic character- istics of the pressure -recording system, but also because of uncertainty regarding the applicability of point pres- sures in relation to total loads under these circumstances, The fact should also be noted that, owing to the inertia of the tail structure, momentary pressure increments would not necessarily result in comparable stress increments. The strain-gage and deflection measurements indicate that the initial effect of the stall may result in up loads of the order of twice those loads experienced immediately prior to stalling. After the initial tail-load increment occurs because of wing stalling, the tail is buffeted repeatedly by the fluctuating downwash in the turbulent wake from the stalled wing. The possibility for resonance between the turbulence frequency and certain natural frequencies of the tail structure exists under this condition. The frequency of the horizontal tail in primary bending was 17— cycles ner second and the frequency of the complete tail in torsion of the fuselage was 10 cycles per second. From tests in the NACA full-scale tunnel, the frequency of the turbulence fluctuations from the stalled wing was 12 KA.CA ARR No. 1J4PO5 found to be 5*5 cycles per second at t>5 miles per hour. If this frequency were a linear function of true airspeed, the range wo^ild be from about 13 to 20 cycles per second in the speed range covered by the pull-up tests and, at some speeds, would coincide with the bending frequency of the tail. The tiirbulence frequencies, however, as shown by the pressure records taken at the tail, were seldom actually uniform for more than 2 or 5 cycles. Moreover, where definite frequencies were detectable, the turbulence frequencies appeared to range from about 10 to 35 cycles per second and to be independent of the speed of flight. This lack of regularity in the turbulence pattern was not unexpected because both the angle of attack of the wing and the position of the tail in the wing wake were rapidly varying with time. In two of the pull-up maneuvers, however, resonance with the tail structure occurred when pressure fluctuations of a frequency close to that of the tail were sustained for several cycles. An example of this condition of resonance is shown by the pull-up recorded in figure lij. where a large periodic build-up in stress occurred as a result of a series of regular pres- s-ore fluctuations. Figure 13 shows a somewhat similar condition at a different airspeed. Both records clearly indicate the mechanism by which excessive tail stresses can be produced when tail buffeting occurs. CONCLUSIONS The results of the present tail-load tests with a modern pursuit airplane show the type and the general magnitudes of loadings encountered on the horizontal ta: of a heavily loaded pursuit airplane in accelerated maneuvers. The survey of critical conditions is not complete, however, because no tests were made in the high-speed and diving-speed ranges. In addition, the measurements that were obtained are less complete and less detailed than ere required to present an accurate quantitative picture of the loads, in particular, the loads immediately after the stall and during tail buf- feting. The need for further investigation of these conditions is indicated. The conclusions to be drawn from the present tests are summarized as follows: \. ARR T 'o. U^05 lj (1) In abrupt pull-ups, the critical horizontal-tail loads were up loads and were substantially proportional to tl e ' aximum normal acceleration. For unstalled pull- ups, extrapolation of the test results shows that a total tail lead of 5?00 pounds would he experienced at an acceleration of 8g. Of this total tail load, about 1000 rounds would he due to the manipul tion of the elevator during the pull-up. (2) In unstalled maneuvers with power on, the span- lc iing on the horizontal tail was unsymmetrical. About 800 pounds more up load -:as carried by the left stabilizer than by the right stabilizer. The magnitude of this dissymmetry was essential. - it of the normal acceleration. With power ."1, bhe dis; ry wa s rr eat] y reduce d . (3) In pull-ups to the stall, an abrupl ' se in the tail load occur". . : diately after the stall of the wi -. Data for 1 M cular airplane tested indicate that load incr - -3 of the order of 100 percent of the load just prior to staliir; e obtained. In stalled pull-up 1a suvers, the tail was buffeted repeatedly : turbulent flow from the stalled win . e possibility of excessive stresses due to resonance in this condition was indicated. Lang ley Memorial Aeronautical Laboratory tional Advisory lee for .-.3ronautics Lang le y Field, V a . NACA ARR No.. L'|F05 Hi- H $? e H t-i |in s O < r r| 1-1 k3 ^ 1-1 K < H R <* CO Js, Fh i-^ H K CO O F Ph PL, H PQ {£ O < *=^: s F=^t i=> CO C CD O I H F-^ (DCh OH H (D fcH W X) - ON ON LP. I PIP. co om-p I I I CO C-rH OJ 0N-O_d/LP LPVO _H; f.P I I I I cm O r\j I NO CO I rH rH t-PLN kn. o-_dr o LP rp CD -d -P CD • c •H W -d in' fi -H OJ N\d/- f-CO-^TH vO .:.-> Vh.no rH rH OJ A) CM rrv LP cm rP-drrH rH N S ,NA _H-ro : cm'-H^' o CD o f- - O 4J -d -p cfl+3 cvS -i r-' o COCO LPO- ^J-LP-d-rH OnO NO ONH Li* '> rH rH OJ CM CM CM O H _j-_H / _H/CM -O CMCO rPNONO CArlO J" CM N-N rH OJ rH CM CM CM ■-NO L rH ( CM LPLPCOVD ^j-CO^D rH C— tO. 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(b) Targets painted on left stabilizer. Figure 3.- Installation for photographing tail deflections NACA ARR No. L4F05 Fig. 4 1 • 1 ** I 1 * * / I 1 1 i I ^ m • ? c ■ ! c 1 1 . ) I { V i o cd H a 2 ° s £ o > a I< ct - 1 2 ■4 u " Z IU O £ I- t o u / / \ \ \ V • "^ ) ^ 1 / s \ M3 N >< Qq ^ ^ ^ 03 s "5" ■P & n o ^ •P c ^ O § -P U C5 -P CO VO 1 ■H i *•< o u ~% 2 > ^ .^ <: S3 m o b s o • U i-l <0 ^ ■H * TJ C a) o <\> rH £ ■H 9 ^0 Sr> a -< a) <^> Pi CO ^ l • Qo a* bO ^ /y c * u o Fig. 6 NACA ARR No. L4F05 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS /.o 4p, Aeff oafsboan/ or/fice Jt 4 4> 4 /X> ^£ r faff- /nbaarS orifice a $,— 5° $ .8 \ 6 r ^ y & y V* /■O ^^rig/it oafboanj orifice .<* *> 4- y, 5^ .? Q O/ ~? .* .6 .8 /.O &£ri#hf /hjboarJ or/fice Figure 6.- Calibration of orifices from full-scale- tunnel tests. NACA ARR No. L4F05 Fig. 7 to y > i- a z> o j 2 o > « _l O < u - Z iU O f < z o — \ * \ \ a // \ \ \ \ \ V \ \ \ \ .N *J o □ \ \ \ \ \ \ \ \ \ I ^ ^ ^ ^ <^ Q I § Qi K ^ o § ^5 S Ci § I V) i i .to 3 c o •H o © f-l %-l ID •O a, •P I u cd 01 4-4 o c o c (0 3 J 6 L/oj^D^/jaa Fig. 8 NACA ARR No. L4F05 "" > § \ c c «)• "0 / / r o <\J 'Vl ? t;' 0) 3 CO CO cu • o. • < r-l s O •H C G CD o3 cj E U, a; .. a. c o CO a. >, / / y r ! 8 00 § 1- *- 2 X O o / I ///# § 97 sojoj joidaji^ ^ O l^ c//7 UMOQ Sap | J* X r^ ^N "1 +J cc3 <*-! l-l ■H bo > u 3 o t- 5 4000 2000 O /ice etertrfion j> Figure 10.- Unsymraetrical spanwiae loading indicated by pressure-orifice measurements. NACA ARR No. L4F05 Fig. U & b «& .-^^ 10 10 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. ^ -"x \l< V. \ — 1 $4? ft) ^"^ ^ £ I IS K ft) Q £ 5 la. 3 — \jftl— - Hi v) ^ ^ 1 o&. -*/ orv f/c ?s C 1 ' .— m ._ it ^. 1 L - ^- '"> s- > ~^j 5»" — z eft S/76oan/ sec Figure 11.- pressure , Time history 6f 180° left turn. 30 inches of mercury at 2450 rpm of pressures on left and right stabilizers. Power on; manifold Note dissymmetry Fig. 12 NACA ARR No. L4F05 NATIONAL ADVISORY COMMITTEE FOR AERONAUT Right ot/f£>ac7/Te/ o/~//~/ces /o 10 § i L&f+ cx/f/boa/yf or/f/ces 2* $ /Ou icsj- s. ^* S ** l r" 1 /?/#/? f /r?6oaf-d or/r/ces\^ \ sV- -*** sj (t to /o ^ - — -^ 5 — :*- -^^' \ V — \ V .If 1!! 24-0 zzo 200 /80 8 /O T/'me, Sec /2 /* /6 /8 20 Figure 12.- Time history of 180° left turn. Power off. Note that pressures on right stabilizer are slightly greater than on left stabilizer. NACA ARR No. L4F05 Fig. 13,14 " pe>o/ &/j ^ pvo/ aft p&j&o/pu/ fy//teu£ jo J0J.U3O 40 UO/JKS2/&OO0 /PLUUO/V Fig. 15a NACA ARR No. L4F05 .0 o -I.O I.O I.O o I.O o 1.0 1.0 o 1.0 o 1.0 o 1.0 Load change, -■488 468 78/ 4se 634 Time, 'sec /.7SO 2. 1 GO 2.SOO 2.511 2533 2.SSO 2.S67 2. $83 2.&OQ 2.GI7 2.633 =^zexso 2.CC7 NATIONAL ADVISORY COMMITTEE FOB AERONAUTICS. Time, Loadchange, sec 2.L83 = ^=2.133 2.7 OO 2.1 n 2.7SO ==~ 2767 2.783 2800 2.817 2.833 s= 2.8S0 distance from £ fuselage, in. Figure 15.- Instantaneous beam diagrams of left stabilizer during a 4.?g pull-up to maximum lift. Run 1 of flight ?1B NACA ARR No. L4F05 Fig. 15b .0 8 5 J; «*- \t T3 s e.o i.o o i.o o i.o o i.o 1.0 l£> 1.0 o 1.0 o 1.0 1.0 o 1.0 o 1.0 o 1.0 2168 /25S /re 683 Load chanqe, . Itk 440 ZO£5 2.9/7 Time, sec 23 6 7 2.883 2.900 2.933 2.950 2.9£7 2-983 3.000 3.017 3.0i3 3 DSD NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. 3.217 3233 3.2 SO 3.267 3.283 3.0Q7 80 40 tef' Distance from 80 <£ -fuselage , in. 40 O Figure 15.- Concluded. Fi g. -16a NACA ARR No. L4F05 Load change. fb To+a/ /oad Load _. change J6 change. Time, sec NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. I.SSO I800 1.8/4- I.8Z9 1.8*3 .857 1,871 I.8B6 1.900 1.917 1.933 I.SSO 1.967 80 K) O 4a eo Distance -from £ -fuselage, in. /20 Right Figure 16.- Instantaneous beam diagrams of stabilizer ob- tained during a 2.4g pull-up to maximum lift. Run 1 of flight 24B. NACA ARR No. L4F05 Fig. I6d Load cftanae, 10 H* Total load CheViqeJI^. .0 I I 4/ Load chanyej /A Time, sec 2.4-33 2.4SO 2.467 2.483 2.SOO 2.SI7 B.533 2SS0 Z.567 2.S83 2.6QO 2.617 2.6 33 2.GSO 120 80 40 40 80 IZO Left Distance from <£ fuselage, in Right Figure 16.- Concluded.' national advisory COMMITTEE FOR AERONAUTICS. Fig. 17a NACA ARR No. L4F05 Total toad Tj me. Load change, chanyejb Load change, sec I.SOO 2JOO 2400 2.500 e.5i7 2.5-33 25 SO 2.S67 2.583 2.600 2.617 2.633 2.6 50 40 40 SO Distance -from £ fuselage, in. 120 night 2.667 Figure 17.- Instantaneous beam diagrams of stabilizer obtained during a 4.2g pull-up to maximum lift. Run 2 of flight 24B. NATIONAL ADVISORY ' COMMITTEE FOR AERONAUTICS. NACA ARR No. L4F05 Fig. 17b L otjc/ change } 7o1til had change, l b Ib*& change, Tirr>«. /b 2.3 00 2.917 2.333 2.3 SO 2.9 67 2.963 3.00O 3.0/7 3.033 3-OSO 3.067 3.083 3.100 120 SO 40 O 40 SO 120 left Distance from £ fuse/age , in ■ Right NATIONAL ADVISORY Figure 17.- Continued, committee for aeronautics. Fig. 17c NACA ARR No. L4F05 Load t ' change, 7isn*j vSec 2.683 2.7CO 2.7/7 2.733 2. 7 'SO 2.767 2.783 2.800 2.814- 2.829 2.84-3 2.857 2.871 2.886 Pis Figure 1 40 O 4C 80 I2D tance from <£ -fusela.ge in. Right NATIONAL ADVISORY 7.- Concluded, committee fob aeronautics. NACA ARR No. L4F05 Fig. 18 ^ ° k\ >3 It •^ ^ ^ k ^1 & 2/0 200\ _ 4 NATIONAL ADVISORY COMMITTEE FOB AERONAUTICS. /■O te /.8 22 2.6 30 34- T//?)£ } sec Figure l8.- Time history of 4.2g pull-up to maximum lift. Run 1 of flight 21B. Fig. 19 NACA ARR No. L4F05 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. 1 J i 1 /» /L I 1 ' 1 \ 1 c Sir S IU *# ^m 5.W o 10 £ n ^ vYv^^^^-— l\ A *!\ h \b \A J \Ki> 1 y-r V Vv / \fl — w \S\J> V WN/ vVu V /50 ^ < # /2 /£ 20 24 2.8 3.2 Figure 19.- Time history of a 2.4g pull-up to maximum lift Run 1 of flight 24B. NACA ARR No. L4F05 Fig. 20 II h NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS. .0 \\ s s ^n^ \ aA n o y "VV) d Mil ^ Ml 1 It 3 f>y 1 VVv A, ,1a nA — s*- 2 ^ \ \Jv\ hn A I* / / \.\ / yf \ \x \l\hl p / w ^ ^» v y 2/0 to W Tt \\zoo