Johns, Slaton L

Pierce, Harold B

English

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I 
I , 
RM L52B04 
NACA 
RESEARCH MEMORAND UM 
GUST-TUNNEL INVESTIGATION OF A DELTA-WING MODE L 
WITH THE LEADING EDGE SWEPT BACK 600 
By Harold B. Pierce and Slaton L. Johns 
Langley Aeronautical Laboratory 
Langley Field, Va. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
April 9, 1952 
https://ntrs.nasa.gov/search.jsp?R=19930086977 2020-06-17T13:49:34+00:00Z

IE NACA RM L52B04 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
GUST-TUNNEL INVESTIGATION OF A DELTA-WING MODEL 
WITH THE LEADING EDGE SWEPT BACK 600 
By Harold B. Pierce and Slaton L. Johns 
SUMMARY 
A gust-tunnel investigation of a delta-wing model with the leading 
edge swept back 600 has indicated that the gust load is greater than 
would be predicted by using a slope of the lift curve determined from 
steady-flow force tests . The use of a slope of a lift curve derived 
from simple aspect-ratio considerations resulted in a good estimate of 
the load. 
INTRODUCTION 
Gust-tunnel investigations of the loads on swept wings (refer-
ences I to 3) have indicated that the slope of the lift curve determined 
from steady-flow force tests is not applicable in a gust but rather 
that one dependent on the slope of the lift curve of an equivalent 
straight wing and on the cosine of the angle of sweep' would be appli-
cable. In addition, the results for the swept wings showed that strip 
theory could be used to estimate the effect of gradual penetration into 
a gust. Since the delta wing might be considered a special form of a 
swept wing, it appeared possible that a slope of the lift curve other 
than that determined from steady-flow force tests would apply in a 
gust. A simple relationship for determining the applicable slope of 
the lift curve, however, was not immediately apparent. Accordingly, an 
investigation has been carried out in the Langley gust tunnel with a 
delta-wing model having the leading edge swept back 600 • The purpose 
of the investigation was to determine the magnitude of the gust load 
and to determine if steady-flow aerodynamic coefficients could be used 
successfully in its prediction. This paper presents the results of the 
investigation. 
2 NACA RM L52B04 
APPARATUS AND TESTS 
A line drawing of the 600 sweptback delta- wing model is shown in 
f igure 1 and a photograph of the model is shown as figure 2 . The char-
acteristics of the model and the test conditions are given in table I. 
The airfoil was an NACA 65-006 section at the root and was tapered to 
an NACA 651- 012 section at a point 1 . 5 inches from the tip. The con-
struction of the model was such that it could be considered a rigid 
body for the tests. An accelerometer was mounted under the forward 
part of the fin at the center of gravity. Two lights were mounted on 
the model for the measurement of pitch in free flight. 
The gust tunnel and its equipment are described in reference 4. 
The gust profiles of the sharp- edge gust and 6.5-chord-gradient gust 
are shown in figure 3 as the ratio of the local gust velocity to the 
average maximum gust velocity as a function of mean chords of penetration. 
The tests consisted of eight flights through the sharp- edge gust 
and seven flights through the gust wi th the 6 . 5- chord gradient distance . 
Measurements of forward velocity, gust velocity, normal-acceleration 
increment, and pitch- angle increment were made for each flight . 
Force tests were conducted in the Langley free-flight tunnel for 
the purpose of determining the free- flight trim condition . A steady-
flow slope of the lift curve was determined from the data and is 
included in table I. 
PRECISION 
The measured quantities are estimated to be accurate within the 
following limits for any test or run: 
Normal-acceleration increment, ~n , 
Forward velocity, feet per second 
Gust velocity, f eet per second 
Pitch-angle increment, degrees •• 
g units ••. 
RESULTS AND DISCUSSION 
. . . . . . . 
+0 . 05 
-+0 . 5 
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The records of all flights were evaluated to obtain histories of 
the normal-acceleration increment and pitch- angle increment during trav-
erse through the gust . Representative test results are shown in fig--
ures 4(a) and 4(b). The maximum acceleration increment was determined 
-~-- - -
• 
NACA RM L52B04 3 
from the time history of each test flight. Since minor variations in 
forward speed and gust velocity are obtained from flight to flight, the 
maximum acceleration increments were corrected to a forward speed of 
88 feet per second and a gust velocity of 10 feet per second on the 
assumption that the acceleration increment is directly proportional to 
forward speed and gust velocity. In addition, the corrected data were 
further reduced to the values for zero pitching motion by the approx-
imate method of reference 1. The values of the maximum acceleration 
increment for each flight as corrected for forward speed and gust veloc-
ity together with the values reduced to zero pitching motion are given 
in table II. The average values for both gust shapes were determined 
from the data and are included in the table. 
Examination of the pitch-increment curves of figure 4 shows that 
the delta-wing model initially p i tches up on entering the gust and soon 
afterward pitches down. The effect of this sequence of pitching nlotion 
is evident when a comparison is made of the effect of pltch on the aver-
age maximum values of acceleration increment in each of the gust shapes. 
(See table II.) In the sharp-edge gust, where the maximum acceleration 
increment is obtained quickly, the initial pitching-up motion increases 
the load a small amount. In the gust with the 6.5- chord gradient dis-
tance, however , the maximum acceleration increment occurs after the model 
pitches down so that the pitching motion decreases the load about 
5 percent. 
The rapid pitching motion of the model as it passed through the 
gusts and the reversal of the effect of pitching motion on the maximum 
accelerations which occurs for the change in gust shape from sharp edge 
to 6.5-chord gradient indicate that the pitching motion of the delta-
wing model is probably quite sensitive to the center-of-gravity position. 
Further investigation will be necessary, however, before quantitative 
predictions can be made of the change in load with center-of-gravity 
position. The initial positive pitching motion of the model is believed 
to be a result of the gradual penetration of the delta wing into the 
gusts. The effect of penetration on the pitching motion in the present 
case, however, does not appear as great as that on the swept wings of 
references 1 to 3. 
Prediction of the gust load was made with the use of. equation (2) 
of reference 1 . The slope of the lift curve was that determined from 
force tests and the unsteady- lift functions were those for infinite 
aspect ratio where the unsteady- lift function for penetration of the 
gust was modified by strip theory to account for the gradual penetra-
tion of the delta wing into the gust. The results of this calculation 
are compared with the experimental results reduced to zero pitching 
motion in table III. The comparison shows that use of the slope of the 
lift curve derived from force - test measurements in the calculations 
results in an underestimation of the gust load of some 20 to 25 percent. 
4 NACA RM L52B04 
In order to investigate the discrepancy between the calculated and 
experimental results, three additional · calculations for the delta wing 
were made with only the slope of the lift curve being changed and, for 
comparison, a calculation was made for a straight wing of the same aspect 
ratio. The lift-curve slopes used in the calculations for the delta wing 
were : (1) a slope of 2.61 per radian obtained from lifting- surface-
theory calculations for a 600 delta wing, (2 ) a slope of 3 . 66 per radian 
predicted from the work of Jones based on Munk's airship theory (refer-
ence 5) for low-aspect-ratio pointed wings, and (3) a slope of 3 . 23 per 
radian obtained from the simple aspect- ratio relation 6A/(A + 2) which 
has been used successfully for determining slopes for the prediction of 
gust loads on straight wings (reference 4). The slope of the lift curve 
of 3 . 23 per radian was also used in the calculation for the straight 
wing together with the unmodified unsteady- lift function for penetration 
into the gust. The results of these four calculations are included in 
table III. 
It is apparent from examination of table III that only two of the 
calculations agreed well with the experimental results . Use of the 
lifting- surface theory underestimated the experimental results by 15 to 
20 percent; whereas use of the slope from airship theory overestimated 
them by about 8 percent. The best agreement was obtained when the lift-
curve slope, derived from the simple formula 6A/(A + 2), was used 
with either unsteady- lift functi on. If the assumption is made that 
other factors such as the choice of the unsteady- lift function are cor-
r ect, the effect of the gradual penetration of the delta wing into the 
gust is indicated to be small. On the basis of these results, there-
fore , it is implied that the gust load on a 600 delta wing can be pre-
dicted by use of the slope of the l i ft curve of 3.23 per radian derived 
from the formula 6A/ (A + 2) even when the effect of the gradual pene-
tration of the delta wing into the gust is disregarded . 
Although the implications of the preceding paragraph are in agree-
ment with the experiment , much doubt still exists as to what the real 
situation might be . The results presented indicated that the loads may 
be predicted equally well by methods having different concepts. It is 
apparent, therefore , that the basic principles are not fully understood 
and additional information will be needed before extension of the results 
of thi s paper can be made for the calculation of the gust load on delta 
wings having apex angles other than 600 • 
NACA RM L52B04 5 
CONCLUDING REMARKS 
A gust-tunnel investigation of a delta- wing model with the leading 
edge swept back 600 has indicated that the use of the slope of the lift 
curve, determined by steady- flow force tests in gust- load calculations, 
underestimated the gust load by 20 to 25 percent . Further study showed 
that a good estimate of the gust load for this wing was obtained through 
the use of a slope of the lift curve derived from the simple aspect-ratio 
relation 6A/(A + 2) . Extension of these results to other delta wings 
with different apex angles , however , does not appear warranted at this 
time. Additional investigation will be necessary to establish the 
importance of the center- of- gravity position on the loads induced by the 
pitching motion of the delta wing in a gust . 
Langley Aeronautical Laboratory 
National Advisory Committee for Aeronautics 
Langley Field, Va . 
REFERENCES 
1. Pierce, Harold B. : Tests of a 45 0 Sweptback-Wing Model in the Langley 
Gust Tunnel. NACA TN 1528, 1948 . 
2. Reisert, Thomas D. : Gust- Tunnel Investigation of a Wing Model with 
Semichord Line Swept Back 300 • NACA TN 1794, 1949 . 
3. Pierce, Harold B. : Gust- Tunnel Investigation of a Wing Model with 
Semichord Line Swept Back 600 • NACA TN 2204, 1950. 
4. Donely, Philip: Summary of Information Relating to Gust Loads on 
Airplanes. NACA Rep . 997, 1950 . (Formerly NACA TN 1976.) 
5. Jones, Robert T. : Properties of Low- Aspect- Ratio Pointed Wings at 
Speeds below and above the Speed of Sound . NACA Rep. 835, 1946. 
(Formerly NACA TN 1032 . ) 
6 NACA RM L52B04 
TABLE I 
CHARACTERISTICS OF THE MODEL AND TEST CONDITIONS 
Weight , Ibs •••• 
Wing area, sq ft • 
Wing loading , Ib/sq ft 
Span, ft 0.. 0 0 
Mean geometric chord measured in 
plane parallel to pl ane of 
symmetry, Area/Span, ft • 
Aspect ratio , A ••••• 
Root chord, ft •••• . 
Slope of the lift curve determined 
by steady- flow force tests, per radian 
Center-of- gr avit y position, percent 
mean geometric chord 
Gust velocity, fps • • • • • . . • • 
Forward velocity, fps 
Pitching moment of inertia , ~1~g~ft2 ' : : •• 
11. 7 
6. 92 
1.69 
4. 00 
1. 73 
2. 33 
3.46 
2.4 
o 
10 
88 
. 0. 286 
NACA RM L52B04 7 
TABLE II 
MAXIMUM VALUES OF ACCELERATION INCREMENT FOR EACH FLIGHT CORRECTED 
FOR FORWARD SPEED AND GUST VELOCITY AND REDUCED TO 
ZERO PITCH 
Gust Flight Corrected Corrected 
shape experimental experimental 
Ll~ lH~ax reduced (g units) to zero pitch 
(g unils) 
Sharp edge 1 1.29 1.25 
2 1. 28 1.26 
3 1.36 1.32 
4 1.25 1.23 
5 1. 25 1.22 
6 1.28 1.26 
7 1.34 1.30 
8 1.35 1.30 
Average 1.30 1.27 
6.5- 1 1.01 1.06 
chord 2 1.03 1.07 
gradient 3 1.06 1.10 
4 1.02 1.09 
5 .98 1.04 
6 1.00 1.07 
7 1.01 1.07 
Average 1.02 1.07 
TABLE III 
COMPARISON OF EXPERIMENT AND CALCULATION 
Cal culated lmmax 
(g units) 
Corrected 
experimental Delta wing Straight wing 
t\nmax reduced 
Gust shape to zero pitch Lift ing- Airship-theory Lift-curve Lift- curve (g uni ts ) Measured lift- surface-theory lift-curve slope slope 
curve slope of lift-curve 
slope ( nA/2 ) (6A/(A + 2» (6A/ (A + 2» 2. 4 slope of 
of 3 . 66 of 3. 23 of 3 . 23 2. 61 
Sharp edge 1.27 0 . 95 1 .02 1.35 1.21 1.27 
. 
6 . S- chor d 1. 07 . 86 . 91 1.15 1 . 06 1.10 gr adient 
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NACA-Langley - 4-e-52 - 350 
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Gust-tunnel Investigation of a Delta-wing Model with the Leading Edge Swept Back 60 Degrees

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