Transonic Flutter Investigation of Models of the Sweptback Wing of a Fighter Airplane

I 1’ GPO PRICE $ CFSTI PRICE(S) $

-

. I -

ff 653 July 65 TRANSONIC FLUTTER INVESTIGATION OF MODELS O F THE SWEPTBACK WING O F A FIGHTEX AIRSLANE By Samuel L. Smith 1 1 1 and Robert W. Boswinkle, Jr.

Langley Aeronautical Laboratory Langley Field, Va.

- I

N66 ( A C C E S S I O N 37116 NU,:IIBER) (THRUI

c - i <CATEGORY1 ( N A S A C E O R TMX O R AD N U M B E R ) I

NATIONAL ADVl SORY COMMITTEE

F O R AERONAUTICS

WASHINGTON April 15, 1958 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESEARCH MEMORANDUM TRANSONIC F'LUTTER INVESTIGATION OF MODELS OF THE SWEPTBACK WING OF A FIGHTER AIRPLANE By Samuel L. Smith I11 and Robert W. Boswinkle, Jr.

SUMMARY A transonic flutter investigation has been made of models of the wing of a current fighter airplane. The models were dynamically and elastically scaled in accordance with criteria which include a flutter safety margin. The wings had an aspect ratio of 3.42 and were swept back 41 .lo along the leading edge and 1 9 . 3 ' along the outer part of the trailing edge. A large trailing-edge fillet extended out to 50 percent of the semispan. The investigation was made in the Langley transonic blowdown tunnel and covered a Mach number range from 0.75 to 1.32.

The flutter boundary was located at simulated altitudes below sea level, the models being flutter free at altitudes above sea level.

However, a region in which the models exhibited large responses to the turbulence of the tunnel stream extended to altitudes above sea level at supersonic Mach numbers. The significance with regard to the air- plane of fie large responses of the models is not known. The flutter boundary shifted to higher altitudes but remained below sea level with the addition of 15-percent-chord leading-edge extensions over the outer 35 percent of the semispan.

-

INTRODUCTION The flutter characteristics of the wing of a current fighter air- plane have been under study. The wing is swept back 4 1 . 1 ' along the 1 9 . 3 ' along the outer part of the trailing edge. A leading edge and large trailing-edge fillet extends out to 50 percent of the semispan.

Calculations indicated that flutter would result at transonic speeds at sea level if the stiffness were reduced only slightly. Experimental data on similar wings (refs. 1 to 4) indicated that possibly a sufficient .

' I stiffness margin existed; however, it was felt that the wing in question was sufficiently different from those Of the references to warrant a * separate experimental study.

The investigation was made in the Langley transonic blowdown tun- nel with models which were dynamically and elastically scaled in accord- The wing ance with criteria which include a flutter safety margin.

spar was cantilever-mounted inboard of the wing root and the tests were made at Mach numbers from 0.75 to 1.32 and at simulated altitudes extending to below sea level. The effect of installing a 15-percent- chord leading-edge extension over the outer 35 percent of the semispan was also investigated.

SYMBOLS b typical wing semichord, ft C local streamwise chord, ft Typical model length 2 length scale factor, Corresponding airplane length ?&-pica1 model mass m mass scale factor, Corresponding airplane mass m' mass of exposed panel, slugs Mach number M dynamic pressure, lb/sq ft S value of y at wing tip t time scale factor, Time required for tunnel airstream to move 1 model chord length Time required for airplane to move 1 airplane chord length T static temperature, R

v velocity, ft/sec

c ....... ...............

. . . . . . . . . . . .

........ NACA RM L S A ~ ~ : : :* : : -

e . . . . . .

Y .........................

-

V reduced velocity based on a representative natural frequency, s distance from wing root measured perpendicular to wing root, ft Y x,y streamwise and spanwise coordinates, respectively, defined

in figure 4

stiffness reduction factor used to provide margin of safety in application of model flutter-test results to the airplane P mass ratio., m'/pv static air density, slugs/cu ft P representative natural frequency, radians/sec Subscripts : A airplane M model MODELS Geometry The models were 3.125-percent-size versions of the wings of a current fighter airplane. The wing models had an aspect ratio of 3.42 and were swept back 4 1 . 1 ' along the leading edge and 1 9 . 3 ' along the outer part of the trailing edge. A large fillet at the trailing edge extended out to 50 percent of the semispan. A sketch of the model is given in figure 1 and some of the more important geometric properties are listed in table I. The fact that the plan-form aspect ratio is twice the exposed-panel aspect ratio (table I) is coincidental.

Because of damage to the models at flutter, six models were required in the investigation. Three models (designated wings 1 to 3) were without leading-edge chord-extensions and were intended to be

identical. The other three models (designated wings 4 to 6) had leading-

edge chord-extensions and were intended to be identical. In addition, the only intended differences between the two sets of models were differences caused by the addition of the leading-edge chord-extensions.

4 to 6

Small differences between models 1 to 3 and also between models .

0 . ... . 0.. . 0 . .* . . 0.. 0 .

. . . .

: - ' . : 0 . . . . .

0 . 0 . .

0 e N A C A R M L S A l ?

0 . e .

0 . 0 . .e. .e I.

0 . 0.. e . . e *

s .

a * b did exist, as evidenced by the measured natural vibration frequencies and node lines (presented in the section entitled "Physical Properties").

* The chord-extensions were over the outer 35 percent of the semi- span and increased the local wing chords by 15 percent.

A model with leading-edge chord-extensions is shown mounted in the fuselage mounting block in figure 2. ( A s shown in figure 2, the wings were painted at intervals along the leading edge to aid in observing the motion of the models during the flutter runs .) The wings without leading-edge chord- extensions had a small amount of positive camber and the leading-edge chord-extensions of models 4 to 6 accentuated the camber.

Scaling The nondimensional mass and stiffness distributions were required to be the sane for the model as for the airplane. The mass and stiff- ness levels for the model were obtained by specifying the scale factors for the fundamental quantities involved: length, mass, and time.

The size of the model was limited by tunnel-wall-interference effects, and on the basis of past experience the length scale factor was chosen to be The mass scale factor was obtained from a requirement that the mass ratio p should be the same for the model as for the airplane, which results in In order to locate the simulated sea level near the middle of the tunnel density range available at a Mach number of 1, the density ratio was chosen to be p p = 1.97. This location of simulated sea level allows M/ A altitudes below sea level to be obtained and flutter margins to be indicated for cases where flutter 3oes not occur above sea level.

The time scale - factor was obtained from a requirement that the

reduced velocity V should be the same for the model as for the air- plane, which results in .

0 . 0.0 . . ...............

0 . . 0 . .

. . . . . . .......

. . . . .

NACA RM Lfs8A15 0 . 0.. 0 . 0 .

& *

t=t) -1 2

f o r the model as f o r the airplane, Since the Mach number i s the same the time scale factor may be written The s t a t i c temperature f o r the airplane i s a function of a l t i t u d e TA However, i n the only, and f o r sea l e v e l it was taken t o be 519' R .

. tunnel during a run, the temperature continually drops as a i r i s expended

from the reservoir and the temperatures obtained a t the various f l u t t e r A study of previous f l u t - points during an investigation a r e different.

t e r data indicated t h a t 4 0 8 ' R w a s near the average value of the s t a t i c temperature t h a t would be expected during the present runs, and t h i s value was used t o obtain the temperature r a t i o used i n the scaling: TM/TA = 0.786.

A l i s t of the pertinent wing and flow quantities and the design It may be noted t h a t the i s given i n table 11.

scale factors used q i s used i n the scale f a c t o r s fer some sf the quantities factor i i s t e d . The factor q has the value 0.76 and occurs because the stiffnesses of the models were made 76 percent of those which would r e s u l t from application of the scale factors as specified (eqs. (l), (2), and ( 3 ) ) . The purpose of reducing the model stiffnesses was t o provide a margin of safety i n the application of the model f l u t t e r - t e s t r e s u l t s t o the airplane. Thus the design reduced velocity f o r t h e but t o t h a t of an airplane model i s equal, not t o t h a t of the airplane, having stiffnesses 76 percent of those of the actual airplane.

The dynamic pressure and Mach number are quantities which a r e If the controllable during a run, whereas the temperature i s not.

dynamic pressure and Mach number are considered t o be fixed and a s t a t i c temperature different from the design value i s obtained, both the density and velocity w i l l be different from the values considered i n the scaling. The density and velocity changes r e s u l t , respectively, mass r a t i o and reduced velocity different from the design i n values of values. However, a combination o f reduced velocity and mass r a t i o which can be expressed i n terms of the dynamic pressure - 2 vM -a

&I

PM i s independent of the temperature, and t h i s combination i s exactly sim- lated i n the runs by the expedient of interpreting the simulated a l t i - tude i n terms of dynamic pressure. Thus, the scale f a c t o r i n t a b l e I1 f o r dynamic pressure i s used t o convert the dynamic pressure f o r t h e airplane a t any a l t i t u d e and Mach number t o the dynamic pressure f o r the model a t the same a l t i t u d e and Mach number. The dynamic pressure f o r the airplane i s assumed t o be t h a t calculated by use of the ICAO standard atmosphere ( r e f . 5 ) . It may be noted t h a t , f o r a given a l t i - tude, q/M2 i s a constant.

The e f f e c t of not having the mass r a t i o and reduced velocity of the models exactly equal t o those of the airplane i s believed t o be negligible i n the present investigation. Experience with a wide variety of f l u t t e r models has indicated t h a t , a t l e a s t within the operational

-

l i m i t s of the tunnel, f l u t t e r a t a given Mach number tends t o occur a t a constant value of dynamic pressure regardless of the individual values of density and velocity.

.

Construction The construction of the models i s indicated i n figure 1. The main spar was made of aluminum alloy, and aluminum-alloy r i b s having U-shaped cross sections were welded t o the main spar. The leading and t r a i l i n g edges were of pine. Balsa was used t o f i l l t h e wing t o contour. Lead weights were placed i n the wing a t various locations and the wings were wrapped with s i l k cloth and painted. Each wing panel was i-nstmented with s t r a i n gages on the main spar near the root. The main spar was clamped inboard of the root, as shown i n figure 1, and thus allowed some r o o t f l e x i b i l i t y . The mounting block shown i n figure 2 was made of aluminum alloy.

Physical Properties The first several natural cantilever frequencies and node l i n e s of I n obtaining the data an each of the s i x wings a r e given i n figure 3 .

electromagnetic shaker was used t o excite each panel separately. The shaker stem acted on the extended wing spars a t the locations indicated by x i n figure 3 and the spars were clamped as indicated i n figure 1.

The positions of the node l i n e s were indicated by s a l t c r y s t a l s sprinkled on t h e wings.

.

The r i g h t panel of model 2, which survived the f l u t t e r t e s t s w a s used t o obtain the f l e x i b i l i t y influence coefficients.

undamaged, b

Influence coefficients were obtained a t 22 stations ( f i g . 4) on the wing

by t h e method described i n reference 6. The influence-coefficient matrix i s given i n t a b l e 111. This matrix has been made symmetrical i n t a b l e IV coefficients symmetric t o the diag- by taking the average of each p a i r of onal. The deviation of the coefficients i n t a b l e I11 from the average values i n t a b l e IV gives some indication of the accuracy of the measure- ments. Only 2.6 percent of the coefficients deviate more than 2 percent, and the greate’st deviation i s 3.6 percent.

The r i g h t panel of model 2 was cut into s t r i p s and the center of gravity, mass, and moment of i n e r t i a about the center of gravity of each s t r i p were measured. The data a r e given i n figure 5. Each s t r i p

was then cut as shown i n figure 4 so t h a t each section corresponded t o

one of the influence coefficient stations. The mass and center of gravity of each section were measured and the values are l i s t e d i n f i g - ure 4. The masses given i n figures 4 and 5 f o r the sections and s t r i p s an allowance f o r the material l o s t i n the saw cuts.

include APPARATUS AND TESTS The investigation was made i n the Langley transonic blowdown tunnel, which has a s l o t t e d test section. The t e s t section i s octagonal i n cross section and measures 26 inches between f l a t s . During the oper- ation of the tunnel, a preselected Mach number i s set by means of a variable o r i f i c e downstream of the t e s t section, and this Mach rxaber i s held approximately constant ( a f t e r the o r i f i c e i s choked) while the stagnation pressure, and thus the density, i s increased. The s t a t i c density range i s approximately 0.001to 0.012 slug per cubic foot, and Mach numbers from subsonic values t o a maximum of about 1.4 may be obtained. Because of the expansion of the a i r i n the reservoir during a run, the stagnation temperature continually decreases, and therefore the test-section velocity i s not uniquely defined by the Mach number.

Additional d e t a i l s of the tunnel are contained i n reference 1. Excel- l e n t agreement between f l u t t e r data obtained i n the tunnel and i n f r e e a i r has been observed ( r e f . 7).

Ln the investigation, each model was cantilever-mounted i n the mounting block shown i n figure 2. The mounting block w a s f i t t e d i n t o a s t i n g i n such a way as t o form a fuselage 3 inches i n diameter which extended upstream i n t o the subsonic flow region of the tunnel. This arrangement prevented the formation of shock waves from the fuselage nose which might r e f l e c t from the tunnel walls onto the model. A sketch of the model mounted on the sting and i n s t a l l e d i n the tunnel i s shown i n figure 6. The sting and model weighed approximately L 290 pounds and the system had a fundamental bending frequency of about 15 cycles per second.

r Wire strain gages were mounted on the wing spars near the root and were oriented so as to indicate model deflections about two different axes. The strain-gage signals, the tunnel stagnation and static pres- sures, and the stagnation temperature were recorded on a recording oscillograph. The strain-gage signals were used to indicate the start of flutter and the flutter frequency. High-speed motion pictures were made during some of the runs.

The wings without leading-edge chord-extensions were tested at The wings with leading-edge chord-extensions zero angle of attack.

were tested at -2O angle of attack in an attempt to reduce the static loads.

RESULTS AND DISCUSSION Presentation of Data The results of the investigation are given in table V(a) for the wings without leading-edge chord-extensions and in table V(b) for the wings with leading-edge chord-extensions. The dynamic pressure at the various test points is plotted as a function of Mach number in figure 7 for the wings without leading-edge chord-extensions and in figure 8 for the wings with leading-edge'chord-extensions. Lines of constant simulated altitude are also indicated in figures 7 and 8 .

Each circle symbol in figures 7 and 8 indicates the point of the start of definite flutter and each square symbol indicates the point run without obtaining of the m a x i m dynamic pressure attained during a A dashed line below a symbol defines a low-damping condition.

flutter.

In the low-damping condition, the strain-gage records and the motion pictures indicated periods of nearly sinusoidal, lowly damped oscilla- tions. The point for the beginning of low damping in each run was indefinite and was somewhat arbitrarily chosen. On the other hand, the point for the beginning of flutter in each run in which flutter was obtained was definite and was characterized by rapidly diverging oscil- lations. The low-damping region is indicated for the wings without leading-edge chord-extensions in figure 7 by dotted shading.

The response frequencies of the wings are indicated near most of the data points in figures 7 and 8 . The response frequency for no- flutter or low-damping points was taken as the predominant oscillation frequency of the models; at flutter, of course, the flutter frequency is listed.

* 3R t The flutter node for both configurations investigated involved bending and torsion of the wing with some rotation in pitch at the wing L root. The rotation in pitch of the wing root was possible because, as previously noted, the main spar was clamped inboard of the root.

A typical oscillograph record showing the strain-gage traces during low damping and flutter is given in figure 9.

Interpretation of Results A s stated in the section entitled "Scaling," the stiffnesses of models were 7 6 percent of the scaled airplane stiff'nesses. The simulated altitudes indicated in figures 7 and 8 are thus to be interpreted as altitudes which, if cleared by the model, could be reached with a 32-percent (1/0.76 = 1.32) margin of safety in stiffness by the airplane.

This statement assumes, of course, that in all other respects the model exactly simulates the airplane.

A n alternate interpretation of the results arises from the fact that for most configurations the dynamic pressure required for flutter varies, to a first approximation, directly with the stiffness level.

Thus, a flutter point obtained with the model indicates that the air- plane will flutter at the same Mach number at a simulated altitude corresponding to a dynamic pressure 32 percent higher than that for the model.

Wings Without Leading-Edge Chord-Extensions The transonic: flutter boundary for the models of the wing without leading-edge chord-extensions is located at altitudes below sea level The dynamic pressure for flutter is indicated to be a mini- (fig. 7).

The low-damping region extends at m at a Mach number of about 0.87.

supersonic Mach numbers to altitudes above sea level. With regard to the airplane, the significance of the low damping obtained with the models is not known. Photographs of the wings without leading-edge chord-extensions after flutter are given in figures lO(a) to 1O(c).

Wings With Leading-Edge Chord-Extensions Because of various data-recording difficulties, the flutter points at the three lowest Mach numbers for the wings with leading-edge exten- sions (fig. 8) are known only to an estimated accuracy of tlOO lb/sq ft for dynamic pressure and kO.03 for Mach number. However, the shape of the transonic flutter boundary is shown to be similar to that for the

wings without leading-edge chord-extensions (fig . 7) . Although the

flutter boundary shifted to h - , l @ ; h e ral,$itudes with the addition of the leading-edge chord-extensions, no flutter was obtained at altitudes ebove sea level.

Low damping preceded the flutter points at the lowest Mach numbers, but the location of these points could not be ascertained and they are omitted in figure 8 and table V(b).

A photograph of one of the wings with leading-edge chord-extensions after flutter is given in figure 10(d).

CONCLUSIONS The transonic flutter characteristics of models of the sweptback wing of a current fighter airplane have been studied in the Langley transonic blowdown tunnel. The models were dynamically and elastically scaled in accordance with criteria which include a flutter safety margin.

The scaling was such that if at a given Mach number a certain altitude is cleared by the model, that Mach number and altitude could be reached with a 32 percent margin of safety in stiffness by the airplane.

The following results were obtained: 1. Although the flutter boundary for the wings without leading- edge chord-extensions was located at altitudes below sea level, a region of lowly damped oscillations that extended to altitudes above sea level was obtained at supersonic Mach numbers.

2. With the addition of 15-percent-chord leading-edge extensions over the outer 35 percent of the semispan, the flutter boundary shifted to higher altitudes but remained below sea level.

Langley Aeronautical Laboratory, National Advisory Committee for Aeronautics, Langley Field, Va., December 20, 1957.

a m mom em e e e em+ e e e e em m e m e e m am me e * e @ e m e

. e e e e m e e m m e e am em e e eem ma

r 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 L33Il3a, 19%.

2. Jones, George W., Jr., and Unangst, John R.: Investigation To Determine Effects of Center-of-Gravity Location on Transonic Flut- ter Characteristics of a 4 3 ' Sweptback Wing. NACA RM L33K30, 19%.

3 . Ruhlin, Charles L.: Experimental Transonic Flutter Characteristics of an Untapered, 4 5 ' Sweptback, Aspect-Ratio-4 Wing. NACA RM L55L22, 19%.

4. Land, Norman S., and Abbott, Frank T., Jr.: Transonic Flutter

Investigation of a Fighter-Airplane Wing Model and Comparison With a Systematic Plan-Form Series. NACA RM L55B16, 1955.

5. Anon.: Standard Atmosphere - Tables and Data for Altitudes to

63,800 Feet. NACA Rep. 1235, 1955. (Supersedes NACA TN 3182.)

6. Jones, George W., Jr., and Young, Lou S., Jr.: Transonic Flutter Investigation of Two 6 4 ' Delta Wings With Simulated Streamwise Rib and Orthogonal Spar Construction. NACA RM ~56127, 1957.

7. Bursnall, William J.: Initial Flutter Tests in the Langley Transonic Blowdown Tunnel and Comparison With Free-Flight F l u t t e r Zeszlts.

KACA IiM L52KL4, 1953.

......

............

. . . . ...

. .e..

0 . 0 . . 0 . .

1 2 NACA RM L58A15 0 . 0 . 0 .

........

TABLE I.- GEOMETRY OF MODELS WITHOUT LEADING-EDGE CHORD-EXTENSIONS

Streamwise airfoil section. tip . . . . . . . . Modified NACA 65~006

Streamwise airfoil section. root . . . . . . . Modified NACA 6511007

Leading-edge sweepback. deg . . . . . . . . . . . . . . . . . . 41.1

Trailing-edge sweepback. deg . . . . . . . . . . . . . . . . . 19.3

Span. ft . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.252

Plan-form area based on extension of panels to model center line. sq ft . . . . . . . .

. . . . . . . . . 0.4562

Plan-form aspect ratio based on extension of

panels to model center line . . . . . . . . . . . . . . . . . . 3.42

Fuselage diameter. ft . . . . . . . . . . . . . . . . . . . . . 0.250

Ekposed-panel span. ft . . . . . . . . . . . . . . . . . . . . 0.498

Exposed-panel area. sq ft . . . . . . . . . . . . . . . . . . . 0.1453

Exposed-panel aspect ratio . . . . . . . . . . . . . . . . . . 1.71

.

TABLE 11.- DESIGN SCALE FACTORS OF PERTINENT WING AND F L O W QUANTITIES

= 1.97; - TM = 0.786; = 0.76

1 TA

Design scale factor

I

Quantity Fundamental quantities : Length . . . . . . . . . . . . .

Mass. . . . . . . . . . . . . . .

-112 . Time . . . . . . . . . . . . . .

t = E) 2 0.03525

Derived quantities : Stream velocity . . . . . . . . .

Stream dynamic pressure . . . . .

Moment of inertia . . . . . . . .

Flexibility influence coefficients Natural vibration frequencies . .

Bending and torsional stiffnesses

I 1..122 x 10-

I I d f rl N of n w r-m m o d N nf n w r-m m o rl N d r l r l r l d r l r l r l r l r l N N N a ....... ...............

0 . 0 e . . . 0 . 0 . 0 .

.......

NACA RM L58A15 : : :' : : - . 0 . 0 .

........................

- t-t- 9'9 N ? ? ? ? ? ?\9 N t ? " ? f \ 9 ? \ 9 ? < ?

N N fu) r \ - + f r l u ) O \ r l O L n N L n P - O r r \ N O \ O D Nu) r l r l N N X n r l rl N N N f rl rl N N N f drl rl N . .

a m r- 0 N . .

W a rl OD rl t- rl . . . . . .

c9 W rl r l r l r l r l d . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

H . . . . . . .

H H . . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . .

. . . . . . . . . .

. . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . .

. . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . .

. . . . . . . . . . . . . . . .

rl g O g N O O . . . . . . . . . . . . . . . .

. . . . . . .

%.< a6-q . . . . . . . . .

. . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . .

?

. . . . . . .

p-,"p--IC\O . . . . . . . . . .

t--*\o\mrc\ Nn+rc\-. . . . . . . . . . . . . . . . .

. . . . . . .

4. . . . . . . . . . .

. .

\9 . . . . . . . . . . . . . . . . . .

. . . . . . .

% % Y R . . . . . . . . . . .

. . . . . . .

rl *" . . . . . . . . . . .

m V I . . . . . . . . . . . . . . . . . . .

. . . . . . .

$32 . . . . . . . . . . . .

. , I . , . . . . . . .

. . . . . . . . . . . .

e . . . . . . . . . . . . . . . . . . . .

N " N F N . . . . . . . . . . . . . . . . . . . .

4 . . . . . . . . . . . . . . . . . . . .

9 . . . . . . . . . . . . . . . . . . . . .

rl . . . . . . . . . . . . . . . . . . . . . .

B . ... ...

0 . 0 . 0 .

. . 0 . . 0 . NACA RM L 5 8 A 1 5

0 . 0 . . . . 0 . .

.. ... . . . .e 0 . . . 0 . . 0 . 0.. ..

TABLE V. - COMPILATION OF TEST RESULTS

Response frequency, Panel behavior* .. 9, v, P, T, cps - ring R u n Point lb/sq ft ft/sec slugs/cu ft oR Left Right Left Right (a) W i n g s without leading-edge chord-extensions -

-

1 1 889.1 0.0048 Q Q 1 . 8 7 7 1,921 427.8 ~, 089.2 L L L . 0 9 5 2,228 .0038 411.8 1 2 Q Q L e 0 9 9 2,795 . , 0 7 0 . 8 -0049 395.1 L L t, 243.0 .0042 L.312 3,264 3 7 3 . 6 t, 2 1 5 . 4 Q Q L . 3 1 7 3,924 .m53 354.5 L L 1 . 0 2 8 31390 ! , 0 0 4 . 7 .0067 397.5 1 4 F L 988.1 .0074 ~.0223,622 389.0 L L ~.211 2,821 L , 1 7 4 . 9 .0041 391 8 3 5 ~.2184,118 .0067 Q Q L , 115.3 349 0 X L L, 128.2 L.155 2,524 .0040 397 * 1 1 6 X L . 1 4 0 3,747 .0067 Q ! , 0 % . 7 356.3 X L ~.0562,860 !,054.1 .0051 414.7 X Q ~ . 0 3 0 3,688 .0078 9 7 3 . 4 371 * 7 X L -755 2,435 783.2 0079 447.8 1 8 X .oog8 9 .758 2,794 755.6 413 * 5 X L .0051 .863 2 , 0 8 6 905 * 0 457.7 X .a70 2,418 F 451.6

906.3 - 0059

L X

. 8 1 5 2,698 835.8 - 0077 437.7

t o F X -822 2 , 9 3 3 .0084 835.1 429.5 L N -898 1,971 .0049 414.0 895.7 2 L 1 Q Q .888 2,205 406.1 877.2 .W57 2 L 2 N 946.6 .0062 Q -979 2 , 7 7 3 3 8 9 * 1 L N .om52 410.2 -938 2 , 6 8 3

931 - 2

L 3 F N .0074 -942 3,154 920.7 397.6 - -

-

(b) W i n g s with leading-edge chord-extensions *Panel-behavior code: F - flutter; L - low damping; Q - maximum q , no flutter; X - paneldamsged; N - no flutter.

?Complete records were not obtained on these runs. The values given are estimates based on available information.

....... ...............

..........

R . . . . . .

NACA RM ~ 3 8 ~ 1 5 : E :. : a

- - - . . . . . .

c 6.76 Fuselage--/ Extended s p e r - q I ~ I , I Spar clamped along this l i n e I Figure 1.- Drawing of model.

Lead weights a r e not indicated. Linear dimensions a r e i n inches.

.

I cu .

B t rn Y) d

i

I " J

4 4w E *

N X X h P U f X

I

i .

Y C .3 x I, V I , .3 c .3 c .3 c c o m n m .A .

b o w I S c 0 m .- V L c 3 L . m m m * 3 6 - t a l * E - E u Y e e .

-P -(v _ _ c4-l .A 0

I

i J cc B o rd t

+

- 1 3 7 4

k I .

0 Definite start of flutter 0 Uiximuri dynamic pressure, no flutter _ _ _ bw-damping condition ... .

. . .... .. . . . Lolr-damping region

Numbers beside data points indicate response frequencies i n cps 4,000

> 3,200

d .)

L

f

(D

k 2,400

0 4 E !

B

1,600 8-7 89 1 .o 1 . 1 1 . 2 103 1 8 4 W h number Figure 7.- Transonic f l u t t e r characteristics of wings without leading- edge chord-extensions.

0 Definite start of f l u t t e r Y I Definite start of f l u t t e r ; location of f l u t t e r point estimated _ - - Low-damping condition Numbers beside data points indloate response frequencies in OPS.

t : P 3,200

>

..

E 2,400 a

i

!$

1,600 07 .e 1 . 0 1 . 1 102 B o h nwnber Figure 8.- Comparison of transonic f l u t t e r characteristics of wings with and without leading-edge chord-extensions. (For runs 14, 15, and 17 t h e accuracy of the data i s less than t h a t f o r the other runs, and although low-damping conditions preceded f l u t t e r , they are not indicated here .)

.

ma a a a a a a om am 0 ma. a am. a 0 Z a m 0 0 . a a * a om a a

NACA R M ~ 5 8 ~ 1 5 : a a a f :a : a * - a o m a m a - 25

.

-

c .

Figure 9.- A typical oscillograph record (run 4, wing 1 ) .

.

am a a a a a a .

N A C A - langley Field, Va.