Lauten, William T , Jr

Mitcham, Grady L

English

NASA Technical Reports Server

RM L51B28 
RESEARCH MEMORANDUM 
NOTE ON FLUTTER OF A 600 DELTA WING ENCOUNTERED AT 
LOW-SUPERSONIC SPEEDS DURING THE FLIGHT OF A 
ROCKET-PROPELLED MODEL 
By William T. Lauten, Jr., and Grady L. Mitcham 
Langley Aeronautical Laboratory 
Langley Field, Va. 
NATIONAL ADVISORY COMM ITTEE 
FOR AERONAUTICS 
WASHINGTON 
May 14, 1951 
https://ntrs.nasa.gov/search.jsp?R=19930086580 2020-06-17T14:05:01+00:00Z
, 
ERRATA 
~A~Wll~ 
NOTE ON FLUTTER OF A 600 DELTA WING ENCOUNTERED AT 
LOW-SUPERSONIC SPEEDS DURING THE FLIGHT OF A 
ROCKET-PROPELLED MODEL 
By William T. Lauten, Jr., and Grady L. Mitcham 
May 14, 1951 
Page 8, table III, last column : The units for the values of inertia 
about center of gravity should be in-lb sec 2 instead of ft-lb sec2 . 
Table IlIon page 9 is obsolete and was inadvertently included. 
Figure 5 should be replac ed by the copy attached because of the poor 
quality of the print that was originally issued. 
1 
... 
NACA RM L51B28 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
NOTE ON FLUTTER OF A 600 DELTA WING ENCOUNTERED AT 
LOW-SUPERSONIC SPEEDS DURING THE FLIGHT OF A 
ROCKET-PROPELLED MODEL 
By William T. Lauten7 Jr. 7 and Grady L. Mitcham 
SUMMARY 
An analysis of the flight time history of a rocket-propelled model 
of a 600 delta-wing airplane configuration, fired for the purpose of 
obtaining zero-lift drag data} indicated wing flutter and subsequent 
failure at low-supersonic Mach numbers. This flutter occurred during 
the unpowered decelerating portion of the flight . The behavior of the 
model during flight is discussed and the mass and stiffness character -
istics of a duplicate wing are presented. 
INTRODUCTION 
During the deceler ating portion of the flight test of a rocket-
propelled model of a 600 delta-wing tailless - airplane configuration an 
apparent wing flutter occurred at a Mach number of 1.11 with subsequent 
structural failure at a Mach number of 0.99 . The flight test was part 
of an investigation of the longitudinal stability and control and drag 
character istics of this delta-wing configuration which is being con-
ducted by the National Advisory Committee for Aeronautics . The primary 
purpose of the test was to obtain zero - lift dr ag data . Previous aero -
dynamic tests are reported in reference 1 . 
In order to obtain structural information which was not obtained 
before the flight but which would be of interest with regard to flutter, 
a duplicate of the wing which failed was constructed for laboratory 
tests . The information thus obtained includes natural frequencies and 
str uctural influence coefficients of the complete semispan wing, the 
mass of sections of the wing associated with the influence coefficients, 
and mass and inertia pr opertie s of streamwise strips of the wing. 
2 NACA RM L5lB28 
Since flutter was considered a possible cause of the structural 
failure and since there is little, if any, theoretical or experimental 
information available concerning the flutter of delta wings in the 
transonic - and low-supersonic-speed range, it is believed that the infor-
mation presented herein will be of interest and may serve in some 
capacity as a guide in future design work. This paper presents the 
structural characteristics of the duplicate wing and a discussion of 
data obtained during the flight of the model. 
APPARATUS AND FLIGHT TESTS 
A three-view drawing of the model used in the flight investigation 
is given in figure 1 and the physical characteristics of the model are 
pre sented in table I . Weight and balance data fo r the model are pre -
sented in table II. Photographs of the model are shown as figures 2 
and 3. The model fuselage was constructed of balsa, plywood, mahogany, 
and aluminum alloy and contained the four-channel telemeter and the 
rocket sustainer motor. The model had a wing of triangular plan form 
with 600 sweepback of the leading edge and an aspect ratio of 2 . 31. The 
wing profile at all spanwi se stations was an NACA 65 (06)-006.5 section 
in the free-stream direction. The vert ical fin of the model was of 
triangular plan form with a leading- edge sweepback of 600 and had the 
same airfoil section as the wing. The wings and vertical fin of the 
model were constructed of mahogany and laminated-pattern pine with an 
aluminum- alloy insert along the chord plane at the trailing edge. 
Longitudinal trim was provided by sealed-gap constant-chord trailing-
edge control surfaces built in with an upward deflection of 0.50 to 
produce approximately zero lift throughout the test speed range. 
The technique of launching and boosting the model to supersonic 
speeds was essentially the same as the technique described in 
reference 1. 
The data from the flight te st were obtained by the use of telemeter, 
photography, radiosonde, Doppler velocimeter r adar, and tracking radar. 
Four channels of information were transmitted and recorded by a telemeter 
system as the model traversed the test speed range. The data recorded 
were time histories of normal and longitudinal acceleration, total 
pre ssure, and base pressure . In addition to the telemetered information, 
records from the two radar units supplied time histories of velocity and 
flight path. Motion-picture cameras also recorded the flight visually. 
The normal accelerometer and the recorder galvanometer s have natural 
frequencies of approximately 100 cycle s per second and are damped to 
about two - thirds critical damping. Only a constant-load calibration was 
, 
i 
I 
--------------
NACA RM L51B28 3 
made of the accelerometer response since high-frequency oscillations 
were not expected. No calibration curves are presented since those 
obtained are not valid at the high frequencies encountered. The type 
of accelerometer and recorder galvanometer used in this investigation 
has been tested in the laboratory from 0 to 250 cycles per second and 
has been found to give a true frequency response throughout this range. 
The amplitude response at 200 cycles per second when the two units are 
used in conjunction is estimated to be about 0.06 of the response at 
zero frequency. Therefore, in determining the magnitude of the normal 
acceleration in the vicinity of 200 cycles per second, the recorded 
amplitude must be multiplied by a factor of about 16. 
GROUND TESTS 
Since flutter was not anticipated during the flight test, flutter 
parameters were not obtained prior to the flight. Consequently, 
following the structural failure of the flight model, a duplicate wing 
was constructed for a ground investigation of its mass and stiffness 
characteristics. 
The quantities determined were the natural frequencies of vibra-
tion, structural influence coefficients of the wing, the mass of sec-
tions of the wing associated with the influence coefficients, and the 
mass, moment of inertia, and the center of gravity of streamwise strips 
of the wing. Their values are given in tables I, III, IV, and V. Fig-
ure 4 is a sketch of the wing which shows the root restraint, points of 
load for influence coeffiCients, streamwise strips, and the sections of 
wing whose mass was determined for use with the structural influence 
coefficients. For the determination of the influence coefficients, the 
wing was loaded by means of a system of wires and pulleys and deflec-
tions were measured with dial gages which could be read directly to 
10-4 inches. The symmetrically placed terms in table V have been aver-
aged to agree with Maxwell's reciprocity theorem. The moments of 
inertia of the streamwise strips were determined by the use of a bifilar 
suspension. 
Although the wing used for the laboratory tests could not be 
expected to be an exact duplicate of the wings tested in flight, the two 
wings were built from the same drawings and it is believed that the 
quantities measured should be in good agreement for the two wings. 
~I 
J 
~-- ------
4 NACA RM L5lB28 
RESULTS AND DISCUSSION 
A study of the telemetered record of the flight} a portion of which 
is shown in figure 5 } shows that a high-frequency oscillation of the 
normal - acceleration trace started in the decelerating portion of the 
flight at a Mach number of 1.11 and continued until structural failure at 
a Mach number of 0 . 99 . This oscillation commenced at a frequency of 
205 cycles per second and decreased to a frequency of 185 cycles per 
second immediately prior to wing failure. Figure 6 shows a plot of 
velocity and Mach number against time for a portion of the flight during 
which the oscillation occurred. This oscillation is believed to have 
been caused by wing vibration since previous experience has shown that 
the normal accelerometer will follow wing vibrations (references 2 and 3). 
Although the limitation of the instrumentation system prevents the accu-
r ate determination of amplitude at high frequencies} it may be noted that 
the amplitude} as recorded} indicates a normal acceleration of approxi-
mat e ly ±0.3g. This quantity is believed to be low by a factor of approxi-
mately 16} as pointed out in the section entitled "Apparatus and Flight 
Tests ." Thus} the oscillating load may have been in the order of ±5g. 
An inspection of the flight time history shows an increase in drag 
at the onset of the oscillations. This phenomenon could be associated 
with either flutter or buffeting. Further study leads to the conclusion 
that it was flutter rather than buffeting. The first factor that leads 
t o this conclusion is that the ratio of the frequency of oscillation of 
the wing tested in flight to the natural torsional frequency of the wing 
tested in the laboratory is 0 .74. This value compares favorably with 
ratios of flutte r f requency to torsional frequency which were obtained in 
the Langley 4.5-foot flutter research tunnel for 450 delta wings 
(unpublished data). Secondly} these oscillations were encountered at 
zero lift} and flight tests of other wings identical in plan form and 
section but not in construction and stiffness (references 1 and 4) gave 
no indication of buffeting at lift coefficients near zero, although 
buffeting was encountered at high lift coefficients . Thirdly} the 
frequency of the oscillation was more than double the first bending 
frequency} while in previous flight tests in which buffeting was 
encountered the frequency of oscillation usually has been at or near the 
first bending f re quency. 
CONCLUDING REMARKS 
An analysis of the flight time history obtained during the deceler-
ating portion of a flight investigation of a model of a delta-wing 
(600 sweepback) tailless-airplane configuration indicated wing flutter 
at a Mach number of 1.11 and subsequent structural failure at a Mach 
number of 0.99. 
J 
NACA RM L51B28 5 
The natural frequencies of vibration, the structural influence 
coefficients of the complete semispan wing, and the mass, moment of 
inertia, and center of gravity of streamwise strips of a duplicate of 
the wing were subsequently determined from laboratory tests. These 
data are presented so that this combination of wing structural charac-
teristics may be a.voided in future designs and these data may be of use 
when a flutter theory is developed for triangular wing plan forms. 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Field, Va. 
REFERENCES 
1. Mitcham, Grady L., Stevens, Joseph E., and Norris, Harry P.: 
Aerodynamic Characteristics and Flying Qualities of a Tailless 
Triangular-Wing Airplane Configuration As Obtained from Flights 
of Rocket-Propelled Models at Transonic and Low Supersonic Speeds. 
NACA RM L9L07, 1950. 
2. Cunningham, H. J., and Lundstrom, R. R.: Description and Analysis 
of a Rocket-Vehicle Experiment on Flutter Involving Wing Defor-
mation and Body Motions. NACA RM L50I29, 1950. 
3. Lauten, William T., Jr., and Nelson, Herbert C.: Results of Two 
Free-Fall Experiments on Flutter of Thin Unswept Wings in the 
Transonic Speed Range. NACA RM L51C08, 1951. 
4. Purser, Paul E. : Notes on Low-Lift Buffeting and Wing Dropping at 
Mach Numbers Near 1. NACA RM L5lA30, 1951. 
6 NACA RM L51B28 
TABLE I 
PHYSICAL CHARACTERISTICS OF THE MODEL OF A DELTA-WING TAILLESS-
AIRPLANE CONFIGURATION 
Wing: 
Area, sq ft (including fuselage intercept). 
Span, ft . . . . . . • 
Aspect ratio . . . . . • . . . 
Mean aerodynamic chord, ft 
Sweepback of leading edge, deg .. 
Dihedral (relative to mean thickness line), deg 
Taper ratio (Tip chord/Root chord) 
Airfoil section . . . . • 
Natural frequencies, cps 
First bending . 
Second bending 
First torsion • 
Vertical tail: 
Area (outside of fuselage), sq ft 
Height (outside of fuselage), ft 
Aspect ratio . . . . . . . . . . 
Sweepback of leading edge., deg 
Taper ratio (Tip chord/Root chord) 
Airfoil section . . . . . . . . . . 
Control surface: 
Type ............... . 
Area (aft of hinge line, one), sq ft 
Span (at trailing edge of wing, one), ft 
Chord (hinge line to trailing edge), ft 
Deflection, deg .......... . 
6.25 
3.80 
2.31 
2 .19 
60 
o 
. . . . . 0 
. NACA 65(06) -006.5 
87 
202 
256 
0.81 
0.97 
2.31 
60 
• . . 0 
NACA 65(06) -006.5 
Plain flap 
0.51 
1. 78 
0.37 
0.5 
-------------- -
NACA RM L5lB28 
TABLE II 
WEIGHT AND BALANCE DATA FOR A MODEL OF A DELTA-WING TAILLESS-
AIRPLANE .CONFIGURATION 
Model with rocket fuel: 
Weight, lb . . . . . . . . . . . . . 
Wing loading, lb/sq ft ..•....•. 
Center-of-gravity position, percent M.A.C. 
Moment of inertia (in pitch), slug-ft2 
Model without rocket fuel: 
Weight, lb . . . . . . . . . . . . . . . . 
Wing loading, lb/sq ft ....... . 
Center-of-gravity position, percent M.A. C. 
Moment of inertia (in pitch), slug-ft2 
7 
148.4 
23 .8 
25.5 
9.37 
123.4 
19.8 
20.0 
8.37 
8 NACA RM L51B28 
TABLE III 
PROPERTIES OF STREAMWISE STRIPS OF WING SHOWN IN FIGURE 4 
Streamwise Spanwise station Center of gr avity Mass Inertia about 
center of gr avity 
strips (in. ) (in. from T.E .) (slugs) (ft -lb sec 2 ) 
I 0 to 1.5 16 . 06 0.0296 0 .134 
II 1.5 to 4.5 13 .94 .047 .167 
III 4.5 to 7.5 10. 94 .031 .076 
IV 7·5 to 10.5 7. 97 .019 . 027 
V 10.5 to 13.5 5 . 37 .012 .007 
VI 13.5 to 17. 8 2 . 94 .005 .001 
2 NACA RM L5lB28 9 
TABLE III 
PROPERTIES OF STREAMWISE STRIPS OF WING SHOWN IN FIGURE 4 
Stre amwi se Spanwise station Center of gravity Mass Inertia about 
center of gravity 
strips (in. ) (in. from T.E.) (slugs) (ft-lb sec2) 
I o to 1.5 16.06 0.0296 0.134 
II 1.5 to 4.5 13.94 .047 .167 
III 4.5 to 7.5 10.94 .031 .076 
IV 7·5 to 10.5 7.97 .019 .027 
V 10.5 to 13.5 5.37 .012 .007 
VI 13.5 to 18.3 2.94 .005 .001 
10 NACA RM L51B28 
TABLE D1 
MASS OF NUMBERED PORTIONS OF WING SHOWN IN FIGURE 4 
Portion Mass (slugs) 
1 0.00820 
2 .01194 
3 .00740 
4 .01235 
5 .00970 
6 .00611 
'7 .00942 
8 .0089'7 
9 .00806 
10 .00658 
11 .00493 
12 .00510 
13 .00518 
14 .00520 
15 .00461 
~ station - 1 Deflectio -at station -
1 42 
2 15 
3 47 
4 5 
5 20 
6 51 
7 2 
8 11 
9 28 
10 55 
11 1 
12 5 
13 13 
14 32 
15 58 
TABLE V 
STRUCTURAL INFLUENCE COEFFICIENTS 
GO-lb load; deflections are in 10-4 incheJ 
2 3 4 5 6 7 8 9 10 
15 47 5 20 51 2 11 28 55 
21 36 13 26 45 5 18 34 54 
36 130 17 59 143 7 30 80 160 
13 17 20 24 33 12 27 38 42 
26 59 24 80 110 12 51 100 142 
45 143 33 110 295 13 63 177 351 
5 7 12 12 13 26 25 25 25 
18 30 27 51 63 25 83 96 107 
34 80 38 100 177 25 96 223 281 
54 160 42 142 351 25 107 281 669 
11 
1 
2 
2 
6 
7 
6 
20 
15 
11 
9 
2 2 6 7 6 20 15 11 9 175 
10 14 18 28 30 44 65 61 54 79 
24 43 36 76 102 43 123 175 192 24 
44 100 45 133 229 36 141 309 473 13 
64 173 49 184 412 27 151 417 911 9 
-
12 13 14 
5 13 32 
10 24 44 
14 43 100 
18 36 45 
28 76 133 
30 102 229 
44 43 36 
65 123 141 
61 175 309 
54 192 473 
79 24 13 
253 152 96 
152 377 327 
96 327 726 
76 295 858 
~ 
15 
58 
64 
173 
49 
184 
412 
27 
151 
417 
911 
9 
76 
295 
858 
2090 
~ (") 
:x=-
~ 
I:-i 
\Jl 
Gj 
r\) 
OJ 
t-' 
t-' 
(j) 
Ii) 
f'-.' 
- 13.64 
J. 26.32 
Elevon hirge 
l ine 
- .l! - - -M.A.G:- - - - - -
-=====:::::=::~~=:::..-_.-~ ')< 
00 0 
Top view 
39.48 -----o-j 
71.22 
13.64-, t-- ----45.59 ----- ---1 
C\J 
to 
L 
Sustainer I 
rocket n Oll e 
I 
o ((:\J 
~ 
C\J 
<i 
C\J 
t----19.10 head orif ic e 
Si de v i ew Front view 
Figure 1.- Three-view drawing of the r ocket- powered flight model. 
(All dimensions are in inches.) 
~ 
,...., 
I\J 
~ 
&; 
~ 
t;-i 
VI 
~ 
CX> 
NAeA RM L5lE28 13 
Figure 2 .- Plan view of flight model. 

NACA RM L5IB28 15 
Figure 3. - Side view of flight model. 
-------- - ---

3 NACA RM L'51B28 
Root 
attachment - __ OL 
2 
4 5 
• 
7 8 
II 12 
II I II III IV 
14 
v 
17 
Points of load application and 
deflection measurement 
Streamwise strip 
VI 
Figure 4.- Schematic dr awing of gr ound- test wing . 

.. Toral PI'¥.ssvl(e ----- -..., 
L "",9/ rvo'lna/ Acc~/el"Qb()/) 
WII'1j' ri:lI/vn ~ I 
.~ ~."...,.".",...~.; '~" 
,,'" . 
"., 
"" .. ' 
j,l' 
\ LJas e ,a~e s.s(/I"e f / 
'-----#ol"ma/ ,4cce/cyQ. t /O/7 _____ ---J 
, . 
, . 
c; 
--- -'--- --- --- ---~
- ... 
~ 
Figure 5.- Portion of telemeter record before and during flutter of flight 
model. 
:\. 
~ 
~ 
&; 
~ 
t"i 
\Jl 
~ 
CD 
t-' 
\0 

5 
0 
Q) 
Ol 
~ 
'H 
., 
:> 
., 
t> 
.r! 
0 
0 
ri 
Q) 
:> 
NACA RM L5lB28 
1400 
1200 1.2 
1000 1.0 
800 .8 
~ 
., 
H 
Q) 
600 1 .6 
..c: 
0 
<tS 
~ 
400 .4 
200 .2 
Beginning 
\ of flutter 
I 
Mach number 
21 
Structural 
o 
o ~ ______ ~ ______ ~ ______ -L ______ ~ ______ ~ ________ ~ ____ ~ 
2.0 2.8 3.2 4.4 
Time, t, sec 
Figure 6.- Plot of velocity and Mach number against time for a portion 
of the rocket-model flight. 
NACA - La ngley Field , Yd. 


Note on Flutter of a 60 Degree Delta Wing Encountered at Low-supersonic Speeds During the Flight of a Rocket-propelled Model

Abstract

Similar works

Full text

Available Versions

NASA Technical Reports Server