Cahen, George L

English

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RESTRICTED RM L53 07 
NACA 
RESEARCH MEMORAND UM 
GUST-TUNNEL INVESTIGATION OF THE EFFECT OF LEADING -EDGE 
SEPARATION ON THE NORMAL ACCELERATIONS E XPERIENCED 
BY A 45 0 SWEPTBACK -WING MODEL IN GUSTS 
By George L. Cahen 
Langley Ae r onautical Laboratory 
Langley Field , Va. 
CLASSIFICATION CANCELLED 
BY AUTHOKITY J. W. CKO LEY 
CHANGE Il673 DATE l2-7-53 T.e.F. 
CLASSIFIED DOCUMENT 
ThIs material contains Information afiectlng the National Defense of the United States within the meaning 
of the espionage laws, Title 18, U.S.C., Sees. 79S and 794, the transmission or revelaUon of which in any 
menner to an unauthorized person Is prohibited by law. 
NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
November 24, 1953 
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NACA RM L53J07 RESTRICTED 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
GUST-TUNNEL INVESTIGATION OF THE EFFECT OF LEADING-EDGE 
SEPARATION ON THE NORMAL ACCELERATIONS EXPERIENCED 
BY A 450 SWEPTBACK-WING MODEL IN GUSTS 
By George L. Cahen 
SUMMARY 
An investigation was conducted in the Langley gust tunnel with a 
450 sweptback-wing model having interchangeable round and sharp leading 
edges in order to determine the effect of leading-edge separation on the 
loads experienced by the model in gusts. 
Leading-edge separation served to increase the gust load. The amount 
of increase apparently depends upon the gust-gradient distance and veloc-
ity but cannot be predicted from the results of these few tests. Attempts 
at correlating the increased load values with the lift curve obtained in 
steady flow yielded no consistent results. Thus) it appears that the load 
increase is due to participation of the leading-edge vortex in the unsteady-
lift phenomena associated with the gust and is not directly related to 
the lift curve obtained in steady flow. 
The results of the tuft studies showed a hysteresis effect under 
unsteady conditions in the flow change between the "no vortex" and "vortex" 
regimes and further indicated that the rate of change in angle of attack 
due to the gust and the extent of penetration into the gust are both 
important in determining whether separation occurs. 
INTRODUCTION 
Leading-edge separation resulting in a leading-edge vortex has been 
observed on sweptback wings in various wind-tunnel studies. In steady 
flow the total lift) as well as the other aerodynamic characteristics of 
a sweptback wing) is affected by leading-edge separation) as has been 
shown in references 1 and 2 ) for example. These studies showed also that 
the angle of attack at which separation first occurs depends on various 
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aerodynamic and geometric parameters of the wing; for example, a sharp 
leading edge is conducive to separation. 
A preliminary investigation conducted in the Langley gust tunnel 
with tufts on three sweptback-wing configurations (ref. 3) indicated 
that a leading-edge vortex was present on a 600 sweptback wing but not 
on a 300 and a 450 sweptback wing during flights through gusts. Inas -
much as a leading-edge vortex occurred on the 450 sweptback wing in wind-
tunnel tests, its absence in the gust flights was ascribed to lag in flow 
development under unsteady conditions. 
In order to determine the effect of leading- edge separation on the 
loads experienced by a wing in gusts, the 450 sweptback wing was fitted 
with interchangeable leading edges, one pair conforming to the original 
airfoil section and the other being wedge shaped in order to induce sep-
aration at a lower angle of attack. This paper presents the results of 
load studies in the Langley gust tunnel and tuft studies in both the 
Langley stability tunnel and gust tunnel for the model with the two 
types of leading edges. 
APPARATUS AND TESTS 
The 450 sweptback-wing model of reference 3 was used for these tests. 
Pert inent characteristics of the model are given in table I. The wing 
was modified in that interchangeable leading edges were fitted to it . 
One set of the leading edges conformed to the original airfoil section 
(NACA 0012 section perpendicular to the 50-percent-chord line), and the 
other set was wedge shaped. Line drawings of the model and the leading 
edges are given in figure 1. 
The upper surface of the left semispan of the wing was fitted with 
cotLon- yarn tufts at intervals of 2 inches . Photographs were made of 
the tufts at each test angle of attack during force tests in the 6 - by 
6 - foot test section of the Langley stability tunnel . These tests were 
made at a tunnel dynamic pressure of 9 pounds per s quare foot through 
an angle-of-attack range from 00 to 300 . 
The Langley gust tunnel and its standard apparatus are described in 
reference 4 . The gust - tunnel tests consisted of fl ights of the model at 
a forward speed of 60 mph with the sharp and round leading edges through 
sharp-edge and 12- foot gradi ent gusts of 5, 10, and 15 ft/sec . The sharp-
edge and gradient gust profiles are shown in figure 2 . Time histories 
of the normal acceleration at the center of gravity of the model and of 
the pitch angle of the model were made for eight flights under each of 
the t welve conditions . Photographs of the tufts were made from above 
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during the gust-tunnel flights by means of high-speed sequence-flash 
photography so that five views of the model were obtained for each flight. 
PRECISION 
The quantities measured in the gust-tunnel tests are estimated to 
be accurate within the following limits: 
Acceleration increment, g units 
Forward speed) ft/sec . 
Gust velocity, ft/sec . 
Pitch-angle increment, deg 
Angle of attack, deg 
to.05 
to., 
-to.l 
-to.l 
to., 
Small variations in launching speed and attitude are apt to occur 
in the gust-tunnel flights and produce variations in the model accelera-
tion increment. Inasmuch as these effects cannot be eliminated by cor-
rections to the data, the average acceleration increment for eight flights 
is presented for each test condition. 
The values of angle of attack measured in the wind tunnel are esti-
mated to be accurate within -to.25°. 
RESULTS 
The records from the gust - tunnel flights were evaluated to obtain 
time histories of normal-acceleration increment and pitch-angle increment 
as functions of penetration into the gust measured in mean chords. These 
histories were utilized in the preparation of figures 3 and 4. 
In figure 3, the values of maximum normal-acceleration increment 
are presented as functions of maximum gust velocity . Each of the values 
represents the average of eight flights which were corrected to "no pitch" 
conditions by multiplying by (1 - ~8 ) , where the pitch-angle increment ~ 
U/V 
and the gust velocity U 
eration increment and V 
are both taken at the point of maximum accel-
is the forward speed. 
Representative histories of normal -acceleration increment and angle 
of attack for the t welve test conditions are shown in figure 4. The 
instantaneous gust velocity, pitch angle, and vertical velocity were used 
in determining the angles of attack so that the values given represent 
the instantaneous angle between the wing chord and the resultant relative-
wind vector. 
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Tuft photographs similar to those presented in figure 5 were used 
to determine whether separation occurred for the various flight condi -
t ions . The lift curves obtained in the wind- t unnel tests for the wing 
with the t wo types of leading edges are shown in f igure 6. 
DISCUSSION 
The results of the load s tudies (fig . 3) indicate that, for all 
gust conditions, the maximum acceleration increments for the sharp-
leading-edge wing were greater than for the round-leading-edge wing. 
The values of normal-acceleration increment obtained for the round-
leading-edge wing are in good agreement with the curves calculated 
according to the usual method of computation (ref. 5). For the sharp-
leading-edge wing, the normal-acceleration increment was computed by 
using several different lift-curve slopes corresponding to different 
points along the lift curve (fig. 6), but no consistent correlation with 
the experimental values could be obtained. Thus, it appears that the 
higher maximum acceleration increments experienced by the sharp-leading-
edge wing were due to participation of the leading-edge separation vortex 
in t he unsteady-lift phenomena associated with the gust and were not 
directly related to the lift curve obtained in steady flow. Since the 
method and the lift-curve slope generally used to calculate gust loads 
for swept wings (ref. 5) are both dependent upon potential-flow concepts, 
they should not be expected to apply for the sharp-leading-edge wing with 
leading-edge separation present. The results of these load studies can 
only be taken to indicate, therefore, that when leading-edge separation 
occurs in a gust, the gust load will be increased. The amount of increase 
apparently depends upon the gust-gradient distance and velocity, as shown 
in figure 3, but cannot be predicted from the results of these few tests. 
The indication by the tufts in figure 5(b) of boundary-layer flow 
parallel to the leading edge of the wing illustrates the presence of 
leading-edge separation. This characteristic tuft behavior along with 
the sudden increase in lift-curve slope {fig. 6 at ~ = 5 .50 for sharp-
leading-edge wing and ~ = 9 .50 for round-leading-edge wing) have been 
defined as signs of leading-edge separation in various wind-tunnel studies, 
for example, references 1 and 2 . For the purposes of these tests, it is 
assumed that the tufts respond almost instantaneously and thus indicate 
a given flow pattern equally well in steady and unsteady flow. 
The solid data points for angle of attack shown in figure 4 indicate 
that separation was present. The conditions of separation or no separa-
tion were obtained from photographs such as those presented in figure 5. 
The positions of the symbols for angle of attack in figure 4, in fact, 
indicate the values of penetration into the gust at which the tuft photo-
graphs were obtained. The normal-accelerat ion-increment histories in 
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figure 4 are sample curves, the peak values of which do not necessarily 
agree with those given in figure 3, which are average values for eight 
flights. These curves are presented so that the relative locations of 
acceleration increment and angle of attack may be compared. 
Examination of figure 4 indicates that separation occurred at lower 
angles of attack for the sharp-leading-edge wing than for the round-
leading-edge wing in both the wind-tunnel and gust-tunnel tests and at 
higher angles of attack in the gust tUnnel than in the wind tunnel for 
both wings, except for the round-leading-edge wing in the 15 ft/sec 
gradient gust. (For the round-leading-edge wing in the gradient-gust 
condition of 15 ft/sec (fig. 4(a)), the angle of attack increased to 
about 9.60 , which was the angle at which separation occurred in steady 
flOW, and remained there for a relatively long period of time; thus, it 
appears reasonable for separation to have occurred in this borderline 
region.) It appears from figure 4 that, for unsteady motion, the angle 
of attack at which separation begins is greater than that for steady 
flow when the angle of attack is increasing and that the angle of attack 
at w~lich separation stops is less than that for steady flow when the 
angle of attack is decreasing. It seems then, as might be expected, 
that the rate of change of angle of attack due to the gust and the degree 
of penetration into the gust (or time) are both important in determining 
whether separation occurs. The present study was not conducted in a 
manner which would permit an analysis of these effects, however. 
Reference 3 indicated that leading-edge separation occurred on the 
600 sweptback wing in the sharp-edge gust of 10 ft/sec and reference 6 
gave an experimental value of maximum acceleration increment for the 600 
sweptback wing which was somewhat higher than that calculated using the 
"cosine law" lift-curve slope. Most of the discrepancy in the result of 
reference 6 was attributed to horizontal-tail effect, but in view of the 
results of the present investigation, it would seem that at least part 
of the discrepancy might have been due to leading-edge separation. Con-
sequently, the statements made in references 5 and 6 that the cosine law 
applies for swept wings in gusts do not appear to apply when leading-
edge separation is well-developed, the effect of leading-edge separation 
being to increase the load. The amount of the increase in load appar-
ently depends upon the degree of leading-edge-vortex development which, 
as mentioned previously, depends upon the gust shape and velocity in addi-
tion to the parameters which affect leading-edge separation in steady flow. 
CONCLUSIONS 
The results of an investigation to determine the effect of leading-
edge separation on the loads experienced by a 450 sweptback-wing model 
in gusts indicated the follOWing conclusions: 
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1. Leading-edge separation increases the gust load to a value higher 
than indicated by the usual method of calculation. 
2 . The amount of increase in load apparently depends upon the gust-
gradient distance and velocity . 
3. The increase of the gust load is apparently due to participation 
of the leading-edge separation vortex in the unsteady-lift phenomena 
associated with the gust and is not directly related to the steady-flow 
lift curve. 
4. There is a hysteresis effect under unsteady conditions in the 
flow change between the "no vortex" and "vortex" regimes. 
5. The rate of change in angle of attack and the extent of penetra-
tion into the gust are both important in determining whether leading-edge 
separation occurs. 
Langley Aeronautical Laboratory, 
National Advisory Committee for Aeronautics, 
Langley Field, Va., September 21, 1953 . 
REFERENCES 
1. Cahill, Jones F . , and Gottlieb, Stanley M.: Low-Speed Aerodynamic 
Characteristics of a Series of Swept Wings Having NACA 65A006 Air-
foil Sections (Revised). NACA RM L50F16, 1950. 
2 . Fitzpatrick, James E ., and Foster, Gerald V.: Static Longitudinal 
Aerodynamic Characteristics of a 520 Sweptback Wing of Aspect 
Ratio 2 .88 at Reynolds Numbers From 2,000,000 to 11,000,000. NACA 
RM LBH25, 1948. 
3. Cahen , George L.: A Preliminary Gust-Tunnel Investigation of Leading-
Edge Separation on Swept Wings. NACA RM L52C 20, 1952 . 
4. Donely, Philip: Summary of Information Relating to Gust Loads on 
Airplanes. NACA Rep. 997, 1950. (Supersedes NACA TN 1976.) 
5. Pierce, Harold B.: Tests of a 450 Sweptback-Wing Model in the Langley 
Gust Tunnel. NACA TN 1528, 1948. 
6. Pierce, Harold B.: Gust -Tunnel Investigation of a Wing Model With 
Semi chord Line Swept Back 600 . NACA TN 2204, 1950 . 
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Weight, lb 
Wing area, sq ft 
Span, ft . 
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TABLE I . - CHARACTERISTICS OF THE MODEL 
Mean chord, parallel to plane of symmetry, ft 
Aspect ratio . 
Root chord, ft 
Tip chord, ft 
Sweep angle at 50-percent - chor d l ine, deg 
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9·75 
6.05 
4.25 
1.475 
2·99 
1.90 
0·95 
45 
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Detachable 
/ead/n.J edge 
~ord 
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wedJe-snaped /ec1Ciin3 edge~ 
Figure 1 .- Pl an view of model and a i rfoil sect i on . 
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separation was indicated by the tufts.) 
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NACA RM L53J07 RESTRICTED 
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for each test condition . 
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Gust-tunnel Investigation of the Effect of Leading Edge Separation on the Normal Accelerations Experienced by a 45 Degree Sweptback-wing Model in Gusts

Gust-tunnel Investigation of the Effect of Leading Edge Separation on the Normal Accelerations Experienced by a 45 Degree Sweptback-wing Model in Gusts

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