Angle, Ellwyn E

English

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RM No. L 7J08 
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RCH MEMORANDUM 
CLAS3 I FICAT ION CAJC2LLED 
AUT,IO.dTY Ii . L. Did D8N C nA;~G E U4 55 
DATE 6- 11 -5 3 T.C. FkASEk , J! 
INITIAL FLIGHT TEST OF THE NACA FR-1-A, A LOW-ACCELERATION 
R OCKET -PROPELLED VEHICLE F OR TRANSONIC 
FLUTTER RESEARCH 
By 
Ellwyn E. Angle 
Langley Memorial Aeronautic'(J Laboratory 
Langley Field, Va. ~" Mt.N ON LOAN FROM THE FILES u 
*TIONAl ADVISORY COMMITTEE FOR AERONAUTI 
IMClL!Y AERONAUTICA'. LAS01ATOiltY 
LANGLEY FIELD, I"f MPTON, VIRGINIA 
,TI . 
NAtiONAL ADVfSO Y COMMITTEE ~o 
NATIONAL ADVISORY ~:fOWAt.TTEE 
FOR AERONAUTICS 
WASHINGTON 
June 29, 1948 
CONFi ' 
https://ntrs.nasa.gov/search.jsp?R=19930085828 2020-06-17T16:00:55+00:00Z

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NACA RM No. L7J08 CONFIDENTIAL ~tI& 
NATIONAL ADVISORY COMMITl'EE FOR /lERONA' 
~ ~~ RESEARCH MEMORANDUM ~~ 
INITIAL FLIGHT 'lEST OF THE NACA FR-l-A, A LOW-ACCELERATION ~(,a 
~ ROCKE'Il-PROPELIED VEHICLE FOR TRANSONIC 
FLUTTER RESEARCH 
By Ellwyn E. Angle 
SUMMARY 
The first of a series of flutter rockets, designated the NACA FR-l-A, 
was successfully launched and flown at the Pilotless Aircraft Research 
Station, Wallops Island, Va. Two identical swept wings were tested; the 
left wing failed at 967 feet per second (659 miles per hour, Mach number 
M = 0.89) and the right wing remained on the model throughout the entire 
flight (maximum Mach number of 1.0). 
The experimental value of failure speed obtained at a Mach number of 
0.89 was 76.2 percent greater than the flutter speed obtained using the 
two-dimensional, incompressible, unswept-wing flutter theory. 
INTRODUCTION 
With present-day piloted and nonpfloted aircraft approaching a Mach 
number of unity, the problem of flutter or aerodynamic instability of 
wings is becoming more and more important. This importance arises in 
the speed range between Mach numbers 0.7 and 1.2 where the values of the 
aerodynamic parameters which affect the flutter speed are practically 
unknown. It is therefore important that experimental data be obtained 
by testing wings with predetermined flutter characteristics in this speed 
range. Such results could be used to assist in the development of a 
transonic flutter theory. 
In order to obtain such data, the Pilotless Aircraft Research 
Division is conducting free-flight 'transonic flutter tests by means of 
low-acceleration rocket-propelled test vehicles. Such tests will yield 
informat ion concerning the flutter speed and flutter frequency. 
The design and flutter analYSis of the wings investigated was made 
with the theory developed by and with the cooperation of the Physical 
Research Division of the Langley Laboratory. 
CONFIDENTIAL 
2 CONFJl)ENTIAL NACA RM No. L7J08 
The first of a series of flutter rockets has been launched and was 
equipped for determining the wing failure speed only. The results of 
this flight are presented herein along with a descrlption of the test 
vehicle. 
fh I 
fh 2 
f t 
fa, 
rna, = 21tfa, 
xa, 
a 
I 
xa, + a + 2 
ra, 
b 
SYMBOLS 
natural flexural frequency (first mode) in cycles per se.cond 
natural flexural frequency (second mode) in cycles per 
second 
natural frequency of torsional vibration in cycles per 
second 
torsional frequency about torsional axis in cycles per 
second 
natural torsional frequency in radians per second 
location of center of gravity behind tors i onal axis in terms 
of half-chord 
location of tors i onal axis of wi ng behind mid chord in terms 
of half-chord 
location of center of gravity behind the 25 percent chord in 
terms of half-chord 
mass radius of gyration referred to torsional axis in tel'rns 
of half-chord 
ratio of mass of cylinder of air of diameter equal to chord 
of wing to mass of wing, both taken for equal lengths 
along span 
half-chord length in feet, measured normal to leading edge 
leadi ng-edge length i n feet 
ratio of leadi ng-edge length to chord length 
CONFIJ)ENTIAL 
NACA RM No. L7JoB CONFIDENTIAL 3 
flutter speed, feet per second, determined from reference 1 
(two-dimensional imcompressible flow theory employing 
first bending torsion mode and density at time of 
failure) 
APPARATUS 
Model 
The NACA FR-l-A is essentially a tailless airplane with four 
laminated spruce wings located at the rear of the body. The two hori-
zontal wings are the flutter test wings, and the vertical surfaces are 
for directional stability and have a flutter speed greater than the 
maximum speed of the model. A tailless configuration was chosen in order 
to keep the weight and drag as small as possible. 
A sketch of the model showing physical dimensions and location of 
the different internal units is shown in figure 1. Two photographs 
showing external appearance are shown as figures 2 and 3. 
This model is propelled by a 10OO-pound thrust rocket engine at an 
average calculated acceleration of 2.8 times gravity. It is launched 
without the aid of a booster and flies with power on for approximately 
15 seconds. The physical characteristics of the model and wings are as 
follows: 
Weight, pounds • • • • • • • • • • • • • • • • • • • • • • • • • • • 246 
Fueelage: 
Length, inches •••••• 
Maximum diameter, inches. 
Horizontal wings: 
Weight, pounds (each) • • • • . ••• 
Area (total), square feet ••••••• 
Area (exposed), square feet. 
Span, feet • • . • 
Aspect ratio • . . , , • • • • . , • • • • • . 
Ai rfoil eect i on (normal to the leading edge), • 
Sweepback angle, degrees •••• 
Loading, pounds per square foot • 
Mean aerodynamic chord, inches 
Taper ratio • • • • • • • • • • • 
CONFIDENTIAL 
. . . • 93 
10.625 
6.00 
7.45 
6.18 
5.25 
. •. 3.7 
• • NACA 6s-0():J 
45 
39.8 
. . . . . 17 
1:1 
J 
4 CONFmENTIAL NACA RM No. L7JoB 
Vertical fins t 
Area (total), square feet. • • • • • • • • 3.40 
Area (exposed), square feet. • • • • • • • • • • • • • • 2.22 
Span, feet •••••••••••••••• • • • • • • • • 2.55 
Aspect ratio • • • • • • • • • • • • • • • • • • • • • • • • • 1.71 
Airfoil section (normal to leading edge) ••••••••• NACA 6~00) 
Sweepback angle, degrees ••••••••••••••••••• 60 
Loading, pounds per square foot • • • • • • • • • • • • • • •• 115 
Mean aerodynamic chord •••• • • • • • • • • • • • • • • •• 16 
Taper ratio • • • • • • • • • • • • • • • • • • • • • • • • •• 1:1 
The flutter parameters obtained from measurements made of the hori-
zontal wings are listed below. 
fhl 
fh2 
ft 
fa. 
a 
rc. 
ra.2 
1 
b 
1 
Ie 
Left wing 
. . . . . . . . . . . . . . . 20.7 
· . . . . . . . . . . . . . . 
· . . . . . . . . . . . . . . . . . 
125 
70.7 
69.0 
-0.2 · . . 
· . . . . . . . . . . . . . . . 
. . 
(standard conditions) 
Instrumentation 
• • 0.06 
0.21 
3.09 
0.5 
26.9 
Right wing 
20.4 
124.5 
69.0 
67.2 
-0.2 
0.06 
0.21 
3.09 
0.5 
26.9 
In order to determine the failure speed of the wings, a simple break-
wire circuit is made by routing a wire through each of the test wings. 
This wire is connected to the power supply of a small radio transmitter 
located in the nose of the model. A continuous signal from the transmitter 
is received at the ground station as long as both wings remain intact and 
is recorded on film with chronograph marks synchronized at the time of 
firing. When either wing fails, the signal abruptly reduces to zero and 
the failure point is known. A portion of the record obtained from the 
flight of the NACA FR-I-A at the time of wing failure is shown in figure 4. 
Launching Technique 
The NACA FR-1-A was launched by means of a zero-length type launching 
rack; that is, the rack does not offer any support or restraint to the 
model after it first starts to move. Because of the inherent character-
istics of the rocket engine to build up thrust slowly, it was necessary to 
CONFIDENTIAL 
--- - - -,------
NACA RM No. L7J08 CONFIDENTIAL 
design a break-link assembly that would retain the model on the launching 
rack until a thrust of approximately 800 pounds was developed. 
Flight-Path Approximation 
Emphasis is placed on the determination of the flight path since it 
is important that the altitude at the time of wing failure be known. 
Calculations based on predetermined aerodynamic characteristics were made 
to show the expected trajectory. The results of these calculations are 
shown in figure 5 as the variation of alt i tude, horizontal distance, 
acceleration, and velocity (Mach number) with time. 
Radar Equipment 
The velocity of the model was obtained by a continuous wave Doppler 
radar unit set behind the launching site ~pprox1mately in line with the 
azimuth of the expected flight path. The record obtained - velocity 
against time - is shown by curve d of figure 5. This velocity is the 
component of the true velocity in line with the radar beam. However, 
the maxi mum error involved in correcting the velocity is small and can 
be neglected. 
Radiosonde 
The atmospheric conditions prevailing at the time of flight and 
pertinent to flutter calculations were deterntned by use of a radiosonde. 
The data are presented in figure 6 as a plot of the velocity of sound 
and denSity against the altitude. 
Cameras 
A E-24 aerial camera operating at 2 frames per second was placed 
normal to the vertical plane of the expected line of flight to obtain a 
record of the flight path and attitude of the model during the first 
2 seconds of flight. 
Two tracking movie cameras, a 35-millimeter and a l&-millimeter, 
were placed about 2000 feet from the launching site wi th the line of 
sight of the cameras normal to the vertical plane of the expected flight 
path at the point where the wIngs were expected to fail. 
CONFIDENTIAL 
6 CONFIDENTIAL NACA RM No. L7J08 
One other 35-millimeter movie camera "las used to record a close-up 
of the actual launching. 
RESULTS AND DISCUSSION 
Launching 
The launching of the NACA Fri-l-A was entirely successful in that 
the break-link assembly released the model at the proper t ime and the 
launching was ~erformed smoothly. In order to aid i n the visualization 
of the take-off, a sequence of pictures made at approximately t-second 
and l-second i ntervals is shown in figures 7 and 8, res~ectively. 
2 
Flight 
After the model was free of the launcher, its attitude changed from 
the initial angle of 630 (launching angle) to an angle of 390 in 2 seconds. 
This change is shown in the sequence of pictures in figure 8. The pre-
flight calculation had indicated that the angle of pitch at this time 
would be 410 • This calculation compares favorably with the actual change 
in attitude angle. 
The flight path deter~ined from the data of figure 8 is shown in 
figure 9 superimposed on the calculated flight path. Since these two 
curves compare closely and since the change in attitude compares closely, 
the remainder of the actual flight path is assumed to be close to the 
calculated trajectory. This assumption is further justified by noting 
the close comparison up to the time of wing failure between the calcu-
lated velocity variation and the velocity obtained by radar, as shown by 
c·~es c and d of figure 5. 
Movies showed that the model traversed a smooth trajectory, made no 
a~preciable deviation from the direction of launching, and developed only 
90 0 of roll during the portion of the flight that the test wings remained 
intact. The evi dence of little roll during the first 2 seconds of flight 
can be seen by referring to figure 8. 
Immediately after wi ng failure the model developed a corkscrew 
fl:l ght path. 
CONFIDENTIAL 
--------------~------------- - - -- ---
NACA RM No. L7JoB CONFJ])ENTIAL 7 
Wing Failure 
A motion picture taken or the rlight shm,s that the lert wing 
fa i led near the root. The time or fa illITe of the left wing is recorded 
by the small radio transm:i tter as 9 .875 seconds after the beginning of 
the fl i ght (f5g . 4). Curve d in f igure 5 indicates that the velocity 
of the model at 9 .875 seconds i s 967 feet per second or 659 miles per 
hour (f~ch number 0.89 ). Th~ s velocity i s the measured failure speed of 
the wing. The flutter s peed calculat ed from reference I for both test 
wi ngs (corrected for an altitude of 2000 rt) is about 550 reet per second 
or 374 miles per hour (Mach number 0.505). These values were obtained 
assuming a wing of zero syreepback, of infin:; te aspect ratio, and traveling 
i n an i ncompressible rluid. (See rererence 1.) It i s seen that this 
experimental value obtained at a Mach numcer of 0.89 is 76.2 percent 
greater than that obtained using the tw~-dlmensional, incompressible, 
unswept theoretical value. 
It is i nteresting to note that, by applying a cosine correction -
an approximate correction for sweepback - to the flut t er speed calculated 
by rererence 1, the experimental value obtained at a Mach number of 0.89 
i s only 24.3 percent greater than the corrected calculated rlutter speed. 
The motion picture also shows that the right wing ren~ined on the 
model throughout t he entire flight , probably due to the change in 
configuration and the disturbed and erratic rlow condi t ions existing 
during the corkscrew flight path followi ng failure of the left wing. The 
model attained a maximum ~~ch number of 1.00. 
The f oregoing data on failure and nonfailure of the wings tested 
indicate that the two-dimensional, incompressible, unswept flutter theory 
i s inadequate in determlning failure s peeds of sweptback f:In:~ te-aspect-
ratio wings i n the transon i c range. Similar trends are noted in the 
freely-fallin~body tests reported in reference 2. More data will be 
necessary (flutter frequency as well as failure speed) to extend the 
flutter results of this t est and to provide a bas i s r or the development 
of a trans onic flutter theory. 
CONCLUSIONS AND REMARKS 
The flutter narameters of t be wings tested were such that the flutter 
s 'need calculated assum:",ng an infinite unswept wing in an i ncompressible 
flm' was 374 m:~les per hour (Mach number 0. 505) . In flight, the left 
wing fa i led to complete destruction at 659 m:~ les per hour (Mach 
num"ber 0. 89 ), and the right wi ng remained on the model thr9ughout the 
fl i ght (maximum Mach number 1.00). It is seen that this experi1r.ental 
8 CONF]J)ENTIAL NACA RM No. L7J08 
value obtained at a Mach number of 0.89 was 76.2 percent greater than 
that obtained using the two-dimens ional, incompressible, unswept 
theoretical value. 
More data (flutter frequency as well as failure speed) will be 
necessary to extend the flutter results of this test and to provide a 
basis for the development of a transonic flutter theory. 
Langley Memorial Aeronautical Le.boratory 
National AdYisory Conn:n.:i.ttee for Aeronautics 
Langley Field, Va. 
REFERENCES 
1. Theodorsen, Theodore, and Garr jck, I. E.: ~er.han ism of Flutter - A 
Theoretical and Exper i mental Investigation ~ the Flutter Problem. 
NACA Rep. No. 685, 1940. 
2. Earmby, J. G., and Clevenson, S. A.: Initial Test in the Transoni c 
Range of Four Flutter Airfo:i..ls Attached to a Freely Fallins Body. 
NACA RM No. L7B27, 1947. 
cOnFIDENTIAL 
CONFIDENTIAL 
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·CONFIDENTIAL 
n9ure j- FR-J-A fluTter test vehicle . 
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CONFIDENTIAL 
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Figure 2.- Side view of the FR-l-A with power plant. 
CONFIDENTIAL 
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Figure 3.- Top view of the FR-I-A showing the flutter test wings .. 
CONFIDENTIAL 
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CONFIDENTIAL 
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Figure 4. - Portio n of telemetered record showing time 
of wing failure. 
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CONFIDENTIAL 
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CONFIDENTIAL 
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CONFIDENTIAL ~ 

CONFIDENTIAL 
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CONFIDENTIAL 
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NACA RM No. L7J08 23 
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NACA RM No. L7J08 25 


Initial Flight Test of the NACA FR-1-A, a Low-acceleration Rocket- Propelled Vehicle for Transonic Flutter Research

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