Document
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..... NASA TM X-598
TECHNICAL MEMORANDUM
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TRANSONIC AND SUPERSONIC FLUTTEI_ TRE]flD INVESTIGATION OF A VARIABLE-SWEEp WING By John C. Stones_lfer and Robert C, Goetz _ _ == _ re= Langley Researcb_C e n_ter Langley Air Force Bas:e_ va.
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NATIONAL AERONAUTICS ANDI _PA_E ADMINISTRATION
WASHI_TON Octobe r 1961
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¢' " , .... .:, 1H _ _ _ o g,'Ip • qu, @'qLU Om e • _ wapm _qP NATIONAL AERONAUTICS AND SPACE AI_INISTRATION TECHNICAL MEMORANDUM X-598 TRANSONIC AND SUPERSONIC FLUTTER TREND INVESTIGATION OF A VARIABLE-SWEEP WING* By John C. Stonesifer and Robert C. Goetz SUMMARY An exploratory flutter trend investigation of a tapered variable- sweep wing design has been made in the Langley transonic blowdown tun- nel and in the Langley 9- by 18-inch supersonic aeroelastlcity tunnel at Mach numbers from about 0.50 to 2.55. Three planforms were tested in order to represent the varlable-sweep wing in three sweep positions.
These planforms had inboard panels which were swept back 60 ° at the leading edge, whereas the outboard panels were swept back 25 ° , 60 ° , and 75 ° at the leading edge. The assumed pivot point was located on the quarter chord of the outboard panel at about 25-percent exposed semlspan of the 60 ° configuration. The models were of simple construc- tion and did not simulate Joint flexibility.
The results at transonic Mach numbers indicate a favorable increase in flutter speeds as the wing sweep angle is increased. At supersonic Mach numbers the 25 ° sweptback wing exhibited the highest flutter speeds.
An appreciable difference between data obtained in the two tunnels is attributed to the different values of mass-density ratio obtained at flutter in the two test facilities.
INTROI_CTION An airplane combining the characteristics of low-speed efficiency and supersonic-cruise ability would be useful in many operations. In general_ however, the dual requirements of low-speed efficiency and supersonic flight are not compatible. In order to accomplish this mission, one would have to compromise the performance of the aircraft or be able to alter its configuration in flight. One method of altera- w tion incorporates variation of the wing sweep by rotating the outer wing panels while the inboard panels remain fixed.
f *Title, Unclassified.
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The use of a variable-sweep configuration, being a new concept, necessitates the need for information that would permit the prediction of its flutter trend characteristics. The purpose of this report is to present the data obtained in an exploratory investigation of the flutter characteristics of a variable-sweep wing in three fixed-sweep positions.
The data were obtained in the Langley transonic blowdown tunnel and in the Langley 9- by 18-inch supersonic aeroelasticity tunnel at Mach numbers from about 0.90 to 2.59. Three planforms were tested in order to represent the varlable-sweep wing in three sweep positions.
L These planforms had inboard panels which were swept back 60 ° at the leading edge, whereas the outboard panels were swept back 25 ° , 60 ° , and 7_ ° at the leading edge. Th_ assumed pivot point was located on the quarter chord of the outboard panel at about 29-percent exposed semispan of the 60 ° configuration. The models were of simple con- structlon and did not simulate Joint flexibility.
It is recognized that reliable flutter information for a given design would require a more realistic model which simulates the design stiffness, mass distribution, and Joint flexibility; h@wever, it is believed that the present data will be useful for preliminary considerations.
SYMBOIS a speed of sound, ft/sec b structural streamwlse semichord at root, ft streamwise semichord at station y, ft flutter frequency, cps ff measured bending frequencies (i = l, 2, 3), cps fh, i measured first torsion frequency, cps exposed semispan, ft mass of one exposed wing panel, slugs m M Mach number dynamic pressure, lb/sq ft q ww mo e(c _e w _ qPo _ w _" w vw w • exposed area of one wing panel, sq ft V stream veloclty, ft/sec Y distance from root chord spanwise, ft A sweepback of leading edge of outboard part of wing, deg m mass-density ratio, nondimensional, air density, slugs/cu ft _k measured first torsion frequency, 2_f_, radians/sec Subscripts: A pertaining to average values for left and right panels of con- structlon A models having the same planform adj value adjusted to what it would be for construction A mln minimum adjusted value of 25 ° configuration MODELS The three planforms shown in figure i were investigated. Each planform had NACA 65A005.5 airfoil sections normal to the leading edge and an inboard part of the wing swept back 60 ° along the leading edge.
The outboard panels of the different planforms were swept back 25 °, 60 °, and 7_ ° along the leading edge.
These three configurations simulate planforms that would result as the outer panels are rotated to vary the sweep angle. The assumed pivot point was located on the quarter chord of the outboard panel at about 25-percent semispan of the 60 ° configuration. The division between the inboard and outboard panels is shown in figure 1. In an airplane with a variable-sweep wing, provision must be made to enclose that part of the wing which disappears as the sweep of the outer panels is increased. Depending on how this is accomplished, the wing trailing edge might not be a straight continuous line near the root as it was for the present models. (See fig. 1.) Therefore, in order to provide some structural consistency for the present models, the structural root chord was arbitrarily made the same (5.55 inches as shown in fig. l) for all three planforms by cutting the root chords at the trailing edges for the 25 ° and 75° swept wings. -- , q_ _w_ • w _ ew _ • w 11w mo eew ..
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The models were machined from solid Formica. In order to obtain flutter throughout the dynamic-pressure range available in the tunnels, it was necessary to reduce the stlffnesses of some of the models. The stiffness reductions were accomplished by cutting a pattern of holes and slots through the outboard panels normal t0the chord plane. The inboard part of the wing was not drilled or slotted for it is believed that for an airplane wlth a varlable-sweep wing this fixed inboard wing area will have to be comparatively stiff to support the sweep mechanism and carry the loads. For models with holes and slots_ the wings were wrapped wlth a layer of sllk which was doped and painted to provide a L leakproof surface. The solid models are referred to herein as construc- tion A, and models with progressively larger holes and longer slots, and J thus progressively reduced stiffnesses, are referred to as construc- tions B, C, D, and E. A photograph of a drilled and slotted semispan model (construction E) is shown In figure 2.
Physical characteristics of all wing panels tested are listed in table I. Both full-span and semlspan sting-mounted models were used in the transonic tests. Wall-mounted semispan models only were used in the supersonic tests. In the transonic tests, although the boundary condi- tions for the full-span and semispan models are dlfferent_ experience with these models and past experience with similar models has shown that the flutter speed is not affected. In many cases, the same model was used for several flutter points before it was damaged or destroyed. A model used in more than one run was checked for structural damage by visual inspection and by comparing natural vibration frequencies of the model obtained before and after each run.
Table I also presents the measured natural vibration frequencies and the ratios of the first and second bending frequencies to the first torsion frequency for the various models investigated. For a given planform, the frequency ratios are shown to be relatively constant as the method of construction Is changed. The natural vibration node lines for the various methods of construction also remained relatively constant for a given planform and are shown In figure 3- For the determination of the natural frequencies and node lines, each model was clamped to a steel bench in such a manner that each wing panel could be considered as canti- levered from the root-clamplng block. (See flg. 2.) An acoustical shaker was used to excite the models. Sand grains sprinkled on the wing surface were used to identify the node lines.
APPARATUS AND TESTS Transonic Tests The tests were conducted in the Langley transonic blowdown tunnel for the Mach number range from 0.50 to 1.30. The transonic blowdown qem wll @ m _0 DW • • • 4W@ tunnel has an octagonal test section with a slot in each corner and tunnel a measures 2_ inches between flats. During operation of the w preselected Mach number is set by means of a variable orifice down- stream of the test section. This Mach number is held approximately constant after the orifice is choked, while the stagnation pressure, and thus the density, is increased. However, the area of the orifice may also be varied during a run as the stagnation pressure is increased or held constant so that various operating paths of Mach number and density may be followed. Both methods of operation were used in the L present investigation and are illustrated in figure 4.
The statlc-denslty range is approximately O.O01 to 0.012 slug per cubic foot. It should be noted that, because of the expansion of the air in the reservoir during a run, the stagnation temperature contin- ually decreases so that the test-section velocity is not uniquely defined by the Mach number. Additional details of the tunnel are contained in reference 1.
The models were mounted on a l_-- inch-diameter sting which formed a fuselage that extendedupstream into the subsonlc-flow region of the tunnel. This arrangement prevented the formation of shock waves from the fuselage nose which might reflect from the tunnel walls onto the model. The 65-pound sting was considered to form a rigid mount for the models since the mass of the complete support system was very large compared with the mass of a model. The fundamental frequency of the support system was approximately 14.5 cycles per second.
An optical system displayed an image of the model on a ground- glass screen during the runs. When flutter was observed, the airflow was quickly stopped in an effort to save themodel from destruction by flutter. " " Supersonic Tests The Langley 9- by 18-inch supersonic aeroelasticity tunnel is a conventional fixed-nozzle blowdown-type wind _unnel exhausting into a vacuum sphere from a pressure reservoir. The nozzles used gave Mach numbers of 1.30, 1.64, 2.00, and 2.55. At each Mach number, the test- section density varies continuously to a controlled maxlmum density and then decreases. Maximum test-sectlon conditions are depicted in the tunnel performance curves shown in figure 5.
For each Mach number in the supersonic tunnel, the test procedure was essentially the same. The sphere to which the tunnel exhausts and the test section were pumped down to a pressure of approximately -__ ...... I u_l _L 2 pounds per square inch absolute. The control valve upstream of the test section was then opened and the test-section density was allowed to increase until flutter was observed or the maximum density was reached.
The models were cantilever-mounted in a mounting block. The mounting block, in turn, was attached to the head of a ram that was used to retract the models through one side of the tunnel in an effort to save the models from destruction after the onset of flutter. The models were viewed through a window in the opposite side of the test L section.
Instrumentation Resistance-type, electrical strain gages were mounted on the sur- face of all models near the root to establish the occurrence of flutter and to indicate the frequency of the flutter oscillation. These gages were oriented so as to indicate as nearly as possible the separate bending and torsional strains of the wings. A multichannel recording oscillograph was employed to record the time history of the straln-gage signals and the tunnel conditions during the runs. High-speed motion pictures furnished a visual record of the model motions.
RESULTS AND DISCUSSION The basic results of this investigation are presented in table II.
The first three columns identify the models as to construction, model number, and full span or semispan. Two full-span models (models l0 and ll) were tested also as semlspan models after the right panels were damaged by flutter. The column labeled "wing-panel behavior" contains a code system defined at the bottom of table II to describe each data point. The low damping behavior indicated by the code letter D is char- acterized by a period of intermittent bursts of slnusoidal oscillations which sometimes obscured the exact start of flutter. Flutter frequencies as obtained Irom the straln-gage traces are given in the column labeled ff. In most cases, the flutter appeared to be of a limited amplitude type.
The flutter and no-flutter points from table II are plotted in fig- V ure 6 in the form of the nondimensional flutter-speed index as a function of Mach number. It should be noted that for a given planform . .... "" .'" . .'" .
i • wqp ¢, • qe w w wp _B got _ • qll_ 911, _ • • Jlq _9 ._. . --_;_;_,,-. : ....
Thus the flutter-speed index is adjusted for variations of the dynamic pressure by the mass and torsional-frequency characteristics of the models.
For a variety of configurations, past experience has shown that the conventional transonic variation of flutter-speed index with Mach number r, is an almost constant value to a point near M = 1 and then, after a slight decrease, increases wlthMach number. Figure 6 indicates that all three planforms exhibit this conventional variation, but to a different degree for each planform. However, the supersonic flutter boundary for each planform exhibits somewhat unusual behavior in that the boundaries obtained in the two wind tunnels do not fair together but show a defi- nite difference at M = 1.3, a Mach number common to both facilities.
Each flutter boundary as shown is composed of two segments. The two segments when considered separately do, however, resemble conventional flutter trends.
It should be noted that, of the several methods of model construc- tion, only construction E models could be fluttered in the supersonic tunnel. A comparison of construction E models with those of other con- structlons showed there was no shift in model center-of-gravity position or noticeable change in frequency ratio or node line characteristics to cause the difference in the data. A further look at other model prop- erties and test conditions indicates that the difference in the data is most likely due to the different values of mass-density ratio obtained at flutter in the two tunnels. A comparison of density ranges of the two tunnels is presented in figure 7. For these densities, the mass- density ratio at flutter for models tested in the transonic tunnel ranged from about 8 to 30 as compared with 36 to 56 in the supersonic tunnel.
Some support for attributing this difference in the data to density difference between the two tunnels is offered by figure 8. In this fig- ure, data from reference 2 are combined with some unpublished test results.
These combined data show the variation of flutter-speed index with Mach number for an aspect-ratio-4, taper-ratio-O.2, 45 ° sweptback wing. As shown, there also is a difference in the data at a Mach number of 1.3.
These data were obtained in the same two tunnels used in the present investigation. In addition to the experimental data, some calculated flutter boundaries for this same wing are presented in figure 8. These calculations (unpublished), based on the method of reference 3 and made by Yates, agree well with the experimental data and support the conclu- sion that a difference such as occurred in the present investigation can result solely from a variation in the density. The data of reference 4 • v v w_ w_ • _ BIW 6_ J_ _ oo : ° • .. .::: I Q also indicate that appreciable scatter in flutter data results if density changes are large.
The flutter boundary curves in figure 9 show the relative flutter susceptibility of the three planforms over the Mach number range of the investigation. In this figure the dynamic pressures for flutter for a given planform of the different constructions, B, C, D, and E, are adjusted to what would result for construction A and then normalized to the minimum adjusted dynamic pressure of the 25 ° configuration.
This adjusted dynamic pressure is obtained from the following simple L relationship between the flutter-speed indices for models of a given planformbut of different construction: V 2 J represents constructions B, (where A represents construction A and C, D, and E). This equation thus yields qadJ = Ah low speeds the 60 ° and 75 ° sweptback wingsrequire dynamic pres- sures for flutter_ respectively, 1.6 and 2.4 times that required for the 25 ° sweptback wing. It may be noted that the transonic dip in dynamic pressure for flutter occurs at progressively higher Mach num- bers and is less severe as the sweepback angle is increased. All three planforms indicate a favorable trend in Mach number effect following the transonic dips. It should be remembered, however, that for the present models, as the sweep angle was increased, certain changes in stiffness (and torsion frequency, see table I) were obtained and this change in stiffness, of course, influences the relative flutter suscep- tibility of the different planforms. If on the variable-sweep wing a different change in stiffness with sweep angle were obtained, as might be expected with a pivot joint, the relative flutter susceptibility would be different.
Also shown in figure 9 are two curves of constant altitude. The actual altitudes simulated by these curves, of course, depend upon the scale factor required to relate the dynamic pressure for the model to the dynamic pressure for the airplane. However, the altitude lines shown with the flutter boundaries emphasize that the relative flutter
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_ ,_ uet • • @@ lip@ 9 • ww susceptibility of the different sweep positions would have to be con- sidered when programlng the wing-sweep changes for a particular flight plan.
CONCLUSIONS The results of a flutter trend investigation in the Langley tran- sonic blowdown tunnel and in the Langley 9- by 18-inch supersonic aero- L elasticity tunnel of very simple models of a tapered variable-sweep wing design indicate the following conclusions: 1. At low speeds the 60 ° and 75 ° sweptback_ings required, respec- tlvely, 1.6 and 2.4 times the dynamic pressure of the 2_ ° sweptback wing for flutter.
2. The transonic dip in dynamic pressure for flutter occurred at progressively higher Mach numbers and was less severe as the sweepback angle was increased.
3. A favorable trend in Mach number effect for all three planforms followed the transonic dips.
4. A difference in the data from the two tunnels was obtained and indicates that careful consideration should be given to the mass-density ratio as a design factor when testing flutter models in different tunnels.
Langley Research Center, National Aeronautics and Space Administration, Langley Air Force Base# Va., July 20_ 1961.
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g¢ _w@ I B _ W@ _S • t _ REFERENCES 1. Unangst, John R._ and Jones_ George W., Jr.: Some Effects of Sweep and Aspect Ratio on the Transonic Flutter Characteristics of a Series of Thin Cantilever Wings Having a Taper Ratio of 0.6. NACA RM LS_I13a_ 1956.
2. Unangst, John R.: Transonic Flutter Characteristics of an Aspect- Ratio-_, _o S_eptback, Taper-Ratlo-0.2 Planform. NASA TM X-136, L 19_9.
3- Yates, E. Carson, Jr.: Calculation of Flutter Characteristics for Finlte-Span Swept or Unswept Wings at Subsonic and Supersonic Speeds by a Modified Strip Analysis. NACA RM LSTLIO, 19_8.
4. Yates, E. Carson, Jr.: Some Effects of Variations in Density and Aerodynamic Parameters on the Calculated Flutter Characteristics of Finite-Span Swept and Unswept Wings at Subsonic and Supersonic Speeds. NASA_X-182, 1960.
TAE_ I.- PHYSICAL CH/LRACTERISTICS OF MOIELS INVESTIGATED i !
Type of Model Configuration I Panel I m, fh, l, fh,2, fh,_ cps 1 fh' I/f_ fh' 2/f_ cps construction I I slugs cps cps f'¢_(_ I I(i , GI4@( I I A = 25o; S = 0.1191 sq ft qll q qll q II _'I a , • • _ C n A Full span Left 0.00276 91 _3 675 I 690 0.13 0.45 .46 314 690 I A Full span Right •00276 90 717 .13 q ,II 41 1 I I Left 26O 519 i 605 .50 C Full span •00252 75 .15 C 76 268 522 I .15 .51 Full span Right .00252 • • _ll "l • 41 I I !
81 .16 C .00269 294. 51o I 665 •58 Semlspan @ • 002_0 62 E 238 460 I 593 .13 .52 Semispan E • 00231 % 225 _47 I 523 .13 .50 Semlspan 5_ .49 E • O0231 99 230 47o I .13 Semlspan !: ii:: .0024-2 58 240 455 I %7 .53 E Semispan .13 24.2 4.60 I E 61 560 .13 -53 Semispan .00239 A = 600; S = 0.1105 sq ft
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120 0.48 Left o. 00290 375 778 920 0.15 9 Full span Q ,m g I1_1 q _ .16 121 .49 Fullspan Rt6ht .00_)0 370 750 890 .16 Left .0o28]_ 100 328 6_5 758 .51 I0 Fullspan .16 I0 • 00281 100 .52 Fullspan Right .530 64.0 765 .16 ii Left 310 600 699 .52 Full span • 00250 97 _-eeql, q 4 ,I i q .16 11 • 0024-8 % 302 600 ..... .50 Full span Right .00266 :,: .......,. , 12 75 260 500 600 -15 .52 Semispan .14 • 00259 74 263 514 6OO .51 13 Semispan @ ,I q ,an ,q _,uq ,t .14 • 52 .OO2_ 69 250 4.80 575 Semlspan Ii t 4 15 Semlspen .00257 7_ 271 500 6_0 .15 A = 750; S = 0.1068 sq ft 0.18 16 Left O. 00287 i_5 791 433 0.55 Fullspan i, 0_0 16 .00287 139 796 451 .17 Fullspan Right i, 050 • 0024.8 lll 6_5 833 375 .17 17 Semlspan 18 i01 8OO .16 •0o259 615 %0 .99 _span 625 .14 .50 19 Semlspan .oo2_,, 76 927 7OO .15 .57 2O Semlspan •00229 79 522 300 21 85 550 69O .15 .55 Semlspan .002_ 300 22 85 590 308 700 .15 •56 Semtspan • 00237 Q .... o_ _ _ • . ............. _.H_ .....
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All dimensions Figure I.- Planforms of models investigated.
are in inches.
L-1302 ,ot-clamping block t ¢4 14q (tt(_ _tqccq :°.,, : ....
U • .. • :'°o Strain gages 4_ 41 _O : '. :o'.
L-60-_87._ Figure 2.- Photograph of a model (construction E) with portion of silk covering removed. k_ • oi oO _ • 6 Q Ji W8 • II ej I
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t p I ---I fh, 2 fa fh, 3 ",--- _ (b) For constructions B, (a) For construction A.
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Figure 3.- Typical node lines for models investigated.
2H _,_- -.. :.. :",." ..''" : : . :._ _ 9QI @ • Q@ •• @ • • _pww vv 4,800 4,400 4,000 3,600 (W O 3,2C0 2,800 2,_00 2, 0_0 1,600
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1,200 80O 4OO 0 ,;, .6 .7 .8 .9 1.0 I.I 1.2 1.3 M Figure 4.- Operating characteristics of the Langley transonic blowdown tunnel.
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q o Po 4,000 q, Ib/sq ft 2,000 p, slugs/cu ft M Figure _.- Performance curves of the Langley 9- by 18-inch supersonic aeroelastlcity tunnel showing maximum test-section conditions obtainable• .6 I I I I I I No Flutter flutter Construction 0 • A V • E ¢ _ cl9 I _Flagged symbol - data obtained in supersonlo aeroelastlclty tunnel- C¢C Unflagged symbol - data obtained in transonlo blowdown tunnel _ 1¢¢4 cllql9 q _ _ qlGg 41 _l,Q q el _Gale I .2
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Figure 6.- Variation of flutter-speed index with Mach number for the three planforms investigated.
kO No Flutter flutter Construction I O O A [] • B - <> @ c V • E Flagged symbol - data obtained in supersonic aeroelastlclty tunnel -- Unflagged symbol- data obtained in transonic blowdown tunnel
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Figure 6.- Continued.
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Figure 7.- Comparison of density ranges in the Langley transonic blowdown tunnel and the supersonic aeroelasticity tunnel.
,,, ,, _ _' _3 0_'I:"'1 ' '" L-1302 o (_Flutter Experimental points for wing tested in transonic blowdown tunnel No flutter .6 Experimental point for wing tested in supersonic aeroelastlclty tunnel.
--Calculated curve using density values from tests in transon£c blowdown tunnel.
-- --Calculated curve using density values from tests in supersonic aeroelasticlty tunnel.
lJ ¸11 ¢41(4 (_C_q -5 _(144 4(((4 q: q ( 4 4 q d 4 d q _J _ Q 4 .4 :,,., :....
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i 0 .8 1.0 1.1 .9 !.2 i. 3 1.4 1.5 M Figure 8.- Comparison of calculated and experimental flutter data for an aspect-ratio-4_ taper- ratio-0.2, h9 ° sweptback wing tested in both the Langley transonic tunnel and the supersonic aeroelasticity tunnel.
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o .8 Io0 I._ J.._ I.o I._ 2.0 2.2 2.l_. 2.6 M I Figure 9.- Flutter boundaries in terms of dynamic pressure applicable to construction A for F_ __N each planform as a function of Mach number with constant altitudes illustrated.
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