Mcfarland, Donald R

Pierce, Harold B

English

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I	 •'X6:64.290 CONFIDENTIAL	 RM L57D22a 
NACA	 FIL 
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RESEARCH MEMORANDUM 
'ciLiA! 2 
EXPERIMENTAL RESULTS ON WING LOADS DUE TO BLASTS

By Harold B. Pierce and Donald R. McFarland 
Langley Aeronautical Laboratory

Langley iField, Va. 
FC.4TIOt4 CI'A!GE!) T ;EO AUT nr;TY /
	 _________________________ 
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ClSlYnaL_ 
This material contains Information affecting the National De*
	
j	 .na within the mea 
of the espionage laws, TItle 18, U.S.C., Sets. 793 and 794, t	 1atIonofwI4teW 
manner to an unauthorized person is prohibited by law. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTiCS 
WASH I NGTOF4 
May 28, 1957 
CON Fl DEN TIAL 
!(Jz	 LillJt)9
https://ntrs.nasa.gov/search.jsp?R=19930089820 2020-06-17T06:37:27+00:00Z
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CONFIDENTIAL NACA RN L57D22a 
	
NATIONAL ADVISORY COMMITTEE FOR Ma0NAUTICS 	 q 
RESELARCH MEMORANDUM 
E)CPERfl1ENTL RESULTS ON WING WADS DUE TO BLASTS 
By Harold B. Pierce and Donald R. McFarland 
SUMMARY 
Recent data obtained in the investigation of the loads caused by 
blast-induced gusts are presented. These data indicate that, when the 
gusts cause angle-of-attack changes to a point well beyond the stall 
angle, a load peak travels along the chord which is caused by a vortex 
formed by the diffraction of the blast wave around the airfoil. When 
the angle caused by the gust is below the stall angle of the airfoil, 
it appears that the loading may be calculated by existing unsteady-lift 
theory for a wide range of the blast position with respect to the wing. 
INTRODUCTION 
This paper is concerned primarily with some recent information on 
the traveling-load peak which was found to occur when an intense blast-
induced gust increases the angle of attack of the wing to a point well 
above the steady-flow stall angle (ref s. 1 and. 2). In addition, some 
data will be presented for the conditions where weak blast waves strike 
from directions normal to the wing surface and nearly parallel to the wing 
span, producing angle-of-attack changes to angles less than the stall angle. 
SYMBOLS 
resultant pressure, measured fran conditions existing prior 
to blast-wave arrival, lb/sq in. 
q	 dynamic pressure, lb/sq in. 
airstream movement, or aerodynamic time, chords 
angle of attack, deg 
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2	 CONFIDENTIAL	 NACA EM L57D22a 
DISCUSSION 
Figure 1 is presented as a review of the information previously 
available on the traveling-load peak. Shown are three sets of load dis-
tributions plotted as load coefficient p/q against percent wing 
chord. These distributions are for three successive times after blast-
wave arrival as represented by the values of 0.53, l.24. , and 2.31 chords 
of travel of the model. The dashed lines with the points represent the 
model flight data, whereas the solid lines were calculated by simple 
potential-flow theory ignoring stall and unsteady-lift effects. Two 
unusual features are apparent in the model flight data. The first is 
that a loading peak was found that traveled rearward as time passed, and 
the second is that this peak showed a larger load than the potential-
flow calculations which would normally be considered an upper limit to 
the loading. 
The traveling-load peak was believed to result from a vortex that 
had been formed when the blast wave diffracted around the leading edge 
of the airfoil and subsequently had been pushed rearward over the upper 
surface by the steady airflow over the wing. However, no direct evidence 
was, available to substantiate this premise. Accordingly, an experimental 
investigation was undertaken in the laboratory to visualize the traveling 
vortices by means of schlieren techniques and to determine how the vortices 
would move along the chord. The manner in which this was accomplished is 
shown schematically in figure 2. At the right, an airfoil having one-
twentieth the chord of the wing of the model used in the free-flight tests 
was placed in the airstream issuing from the slot in the flat surface. 
The airfoil was struck by the blast wave from the actual explosion at the 
left. The airstream velocity and the gust velocity induced by the blast 
wave were set at the values obtained for the free-flying model, but the 
duration of the blast wave was one-twentieth of that to which the free-
flying model was subjected in order to account for the reduction in the 
chord length of the airfoil. 
A series of tests was made with the schlieren apparatus adjusted to 
take photographs of the airfoil and flow at various time intervals after 
the blast wave struck. A sequence of three photographs, with the air-
foil oriented the same as shown in figure 2, is given in figure 3. The 
picture at the left, which was taken shortly after the blast-wave front 
passed, shows that vortices have actually been formed by the diffraction 
of the blast wave around the airfoil and that the one formed at the 
leading edge has moved slightly back along the chord.. The second and 
third pictures, taken at later times, show that this vortex has moved 
still farther back along the chord and successively ,
 become less distinct 
and more spread out. The reduction in distinctness is an indication of 
loss of intensity and conforms to the results from the flight tests shown 
in figure 1 where the loading peak reduces and spreads out as it moves 
along the chord.
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3 COI'IFIDEWrIAL NACA RN L57D22a 
In reference 1, it was found that the load peak traveled along the 
chord at a velocity slower than that of the airstream. This is illustrated 
in figure 1 by the fact that the load peak is still on the airfoil in the 
third load distribution shown although the airstream has moved more than 
two chords since the blast wave struck. The schlieren photographs of 
figure 3 also show the same thing for the vortex. For example, in the 
middle photograph in figure 3, the vortex which was formed at the 
leading edge is seen to be only about a third of the way along the chord 
although the airstream has moved more than 0.9 of a chord. Note that 
the vortex formed at the trailing edge appears to move at airstream 
velocity, for it is nearly 1 chord away from the airfoil. The movement 
of the load peak and the vortex are compared in figure 1. The positions 
of each along their respective airfoils are shown in terms of percent 
chord plotted against the chord-length movement of the airstream across 
the two airfoils. The circles represent the load peak and the squares, 
the vortex. It is obvious that, despite the twenty-to-one difference in 
scale, the movement of the load peak and the vortex is the same function 
of chord-length travel of the airstreazn. The average velocity of travel 
of the load peak and vortex appears to be about one-third stream velocity 
since, in the time they take to pass over the airfoil, the airstream has 
moved nearly 3 chords. 
Recent experimental results are now considered from another phase 
of the investigation, that of determining the loadings when a weak blast-
induced gust strikes a wing and causes an angle change less than the 
stall angle. Earlier unpublished results have indicated that when a 
weak blast wave strikes normal to the wing, the loading can be calculated 
by using existing unsteady-lift theory. Questions have arisen as to 
whether this procedure would apply if the blast wave struck fran a direc-
tion well to the side, say at an angle of 20° to 30° to the wing surface. 
Figure 7 presents the load distribution along the wing chord from one 
test for such a blast direction together with the results previously men-
tioned for the blast wave striking normal to the wing. Both represent 
the load distribution existing at the same very short time after the blast 
waves struck. Again, the data are plotted as load coefficient p/q as 
a function of percent of wing chord. The circles are the experimental 
points associated with the orientation at the right where the blast wave 
strikes normal to the wing, and the squares are the results obtained when 
the model rolled as shown so that the blast wave struck at an angle of 
about 28° to the wing surface. The blast wave and the induced gust 
velocities were the same for each condition. In making the unsteady-
lift calculations, the full gust velocity imposed by the blast wave was 
used in determining the distribution shown by the solid line for the 
condition where the blast wave struck normal to the wing. For the rolled 
condition, however, a simple resolution of vectors was used to determine 
the component of the gust velocity which acted normal to the wing, and 
the result of that unsteady-lift calculation is plotted as the dashed 
line. Both load-distribution calculations are seen to agree well with 
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NACA RM L57D22a 
the experimental data along most of the chord. Near the leading edge, 
the circle representing the flight result for the blast normal to the wing 
is somewhat lower than the potential-flow calculation, as often happens, 
but the square representing the rolled condition is well above its corre-
sponding calculated value. The' flow phenomena which caused this are not 
understood, and, unfortunately, the pressure cell at the 12-percent-chord 
station did not oparate on this flight so it could not be determined if 
the effect was confined only to the leading edge. The square at the 
28-percent-chord station is also high, and again it is not known whether. 
it represents some carryover from the condition found at the leading 
edge or if its position reflects instrument malfunction. It appears, 
however, that by a simple resolution of the gust velocity vectors, the 
loading over the major portion of the chord can be calculated over a 
wide range of blast incidence angles. 
SU1MARY OF RESULTS 
Recent data obtained in the investigation of the loads produced by 
blast-induced gusts have shown that: 
1. When the angle-of-attack change caused by the gust is much 
greater than the stall angle of the airfoil, the load peak traveling 
along the chord is caused by a vortex formed by the diffraction of the 
blast wave around the leading edge and its movement is a function of 
aerodynamic time or chord-length travel. 
2. For weak blast waves causing small angle-of-attack changes, it 
appears that the chord loading may be calculated by existing unsteady-
lift theory. for a wide range of the blast position with respect to the 
wing by making a simple resolution of the gust velocity vector to a 
plane normal to the wing. 
Langley Aeronautical Laboratory, 
National Advisory Committee for Aeronautics,

Langley Field, Va., March 5, 1957. 
C ONFIDEWI'IAL
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NACA RN L57D22a	 CONFIDENTIAL	 5 
REFERENCES 
1. Pierce, Harold B., and Reisert, Thomas D.: Initial Experimental 
Investigation of the Aerodynamic Load on the Wing of a Model Caused 
by a Blast-Induced Gust That Increases the Angle of Attack Into the 
Stall Region. NACA RN L55H22b , 1955. 
2. Pierce, Harold B., and. Spahi, Raymond J.: Experimental Investigation 
To Determine the Loads on a Horizontal Tail of a Model Caused by a 
Blast-Induced Gust. NACA RN L57A28a, 1957. 
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EXPERI M ENTAL AND CALGU LATED LOAD DISTRIBUTIONS 
	
-8-	
r0.53 CHORDS	
FLIGHT 
CALCULATIONS 
	
O	
28°	 -	 POTENTIAL-FLOW 
0 
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.24 CHORDS	 T 2.31 CHORDS 
aI2° 
: o22 _ 
20 40 60 80 1000 20 40 60 80 100 
%WING CHORD 
Figure 1 
ARRANGEMENT FOR VORTEX-FLOW STUDY 
BLAST 
WAVE 
EXPLOSION
FLAT
FIELD OF 
SCHLIEREN-
SUPP0RT- rT+-
Figure 2 
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VORTICES FROM BLAST-WAVE DIFFRACTION 
	
r=O.28 CHORDS	 r=O.93 CHORDS	 t-=I.31 CHORDS 
L-57 -178 
Figure 
LOAD-PEAK AND VORTEX TRAVEL ALONG CHORD 
	
4-	 0 LOAD PEAK 
	
I	 0 VORTEX 
3
0 
AIRSTREAM	 0 
	
MOVEMENT, 2	 00 
i-,CHORDS	 0	
0 
0 00 
0 
0 
00 
_0 0 U	 i	 I 
	
0	 20	 40	 60	 80 
DISTANCE TRAVELED, % CHORD 
Figure i. 
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TIAL	 NACA FIN L57D22a 
EFFECT OF BLAST - GUST DIRECTION ON LOAD DISTRIBUTION 
MODEL I FLIGHT CALCULATED ORIENTATION I 
0	 20	 40	 60	 80	 100
% WING CHORD 
FIgure 5 
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NACA - Langley Field, Va.
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CON Fl DENTIAL


Experimental results on wing loads due to blasts

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