Cooper, George E

Rathert, George A , Jr

English

NASA Technical Reports Server

" 
RM No A8c2 
NACA 
RESEARCH MEMORANDUM 
VISUAL OBSERVATIONS OF THE SHOCK WAVE 
IN FLIGHT 
By George E. Cooper and George A. Rathert, Jr. 
Ames Aeronautical Laboratory 
Moffett Field, Calif. 
--Ii l 
NA llONAl ADVISORY COMMITTEE 
FOR AERONAUTICS 
WASHINGTON 
May 24, 1948 
https://ntrs.nasa.gov/search.jsp?R=19930085360 2020-06-17T14:24:45+00:00Z

NACA RM No. A8C25 
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
RESEARCH MEMORANDUM 
VISUAL OBSERVATIONS OF THE SHOCK WAVE 
IN FLIGHT 
By George E. Cooper and George A. RathertJ Jr. 
SUMMARY 
This report presents the results of pilot observations and 
photographs of the compression shock wave on the wing of an air-
plane in flight. A detailed description of the test conditions 
necessary and the procedure to be followed in producing the visible 
shock are presented. The physical phenomenon occurring is dis-
cussed~ andJ although the observations presented have been made on 
but three airplanes~ it is indlcated that it should be possible to 
obtain photographs of shocks on any airplane if the correct test 
conditions are duplicated. The pilot's observations of the occur-
rence of an oscillating shock wave and its possible relationship 
with the airplane buffeting are also presented. It is suggested 
that future extension of the method. to the study of shock waves 
over other surfaces may prove fruitful. 
INTRODUCTION 
During high-speed dive tests of a single-engine fighter-type 
airplane, the pilot noticed a visible indication of the normal 
shock wave over the wing of the type that was first reported by 
Major Frederick A. Borsodi of the Army Air Forces. Subsequent 
attempts to duplicate the conditions necessary to the formation 
of this visible shock-wave image met with such success that. an 
analysis of the physical phenomenon involved was made and is 
presented in this report. Certain simple rules were established 
which must be adhered to in producing these visible shock-wave 
images at will. The installation of either l6-millimeter or 
35-millimeter motion picture cameras in three different airplanes 
produced numerous photographs of the shock wave on both the wing 
and the canopy. 
The pilot's observations are used freely throughout the 
2 NACA RM No. A8c25 
discussion, and several possible applications of the method are 
suggested which are based on the pilot's opinion of the method's 
utility as a regular testing technique. 
TEST CONDITIONS AND PROCEDURE 
Physical Explanation of Image 
The conditions necessary to produce visible shock-wave images 
are basically the same as those required for any shadowgraph image 
of a pressure discontinuity. A light source of suitable intensity is 
used to produce parallel light, which is refracted upon passing 
through a varying density medium before falling upon a screen or 
viewing surface. A camera may be added for recording purposes but 
is not essential for viewing purposes. In this particular applica-
tion of the shadowgraph principle, the sun is used as the light source 
and the wing surface as the viewing screen. The types of image pro-
duced are shown in figure 1. Note that the image of the canopy shock 
appears in profile directly above the shadow of the canopy itself, 
while t~e wing shock is in plan form. 
An interpretation of the physical phenomenon involved in the 
wing image can be made by referring to figure 2. Parallel light 
from the sun passes through the low-density air ahead of the shock 
wave and is refracted toward the rear upon entering the high-density 
region behind the shock wave. As the pressure change is greater 
near the surface, the amount of refraction is greater there also. 
Consequently, one would expect a dark band immediately behind the 
shock-wave wing intersection, followed by a light band. An examina-
tion of figures 1 and 2 shows this to be the case. The image of the 
canopy shock shown at the left of figure 1 has a greater width of 
dark and light bands than the wing shock, because of the conical 
shape of the shock wave and its greater distance from the wing which 
served as the background or viewing surface. 
Orientation of Sun and Airplane 
Experiments have shown that the clearest shock-wave images are 
produced when the sun is near its zenith and lying approximately in 
the plane formed by the lateral and directional axes of the airplane. 
A drawing showing the proper orientation of the sun with respect to 
the airplane is presented in figure 3. Visibility of the wing shock 
wave from the cockpit was about the same looking either into or away 
from the sun in the cases where the sun was not at its zenith. 
NACA RM No. A8c25 3 
Canopy shock waves, of course, are produced only on the down sun 
wing and require the sun to be near but not at its zenith. The image, 
in this case, is seen on the inboard portion of the wing where visi-
bility is best from ~he cockpit. 
Considerable difficulty has been experienced in obtaining 
prominent shock-wave pictures during the winter months (Oct. - Feb.) 
due to the low angle of the sun at its zenith. Conversely, during 
the summer months no difficulty was experienced in obtaining strongly 
visible shock waves and, furthermore, the latitude with respect to 
the sun-airplane relationship is the greatest. Satisfactory results 
were obtained during this period with the SUD as much as 200 forward 
of the lateral and directional axis plane as indicated in figure 3. 
An account must be taken of the airplane attitude required to 
obtain the critical Mach number necessary for shock formation. Air-
planes which require a relatively steep angle of dive must be dived 
in some degree toward the SUD; whereas jet aircraft attaining criti-
cal Mach numbers in level flight may be expected to obtain the best 
results while the sun is directly overhead as is indicated from a 
study of figure 3. 
Location of Pilot and Camera 
Location of the pilot with respect to the shock-wave image on 
the wing is also of great importance in relation to its visibility, 
and is probably the outstanding reason why observation of the image 
has been reported only once before. 
It is interesting to note that the pilot was flying the same 
model of airplane when he first noticed the shock wave as was 
Major Borsodi when he made his initial observations. 
Subsequent flights in other airplanes, where a visible shock 
wave was produced, indicate that this phenomenon is not confined to a 
single airplane, however, but has been clearly observed on two other 
types. Additional corroboration is found in reference 1, where shock 
waves were observed at the wing-tip tank intersection. 
On the initial airplane, the pilot was located almost squarely 
opposite the wing-chord station near where the shock occurs. If 
the line of the shock wave were extended on this airplane from one 
wing to the other, it would pass almost directly through the pilot's 
seat. Photographs were obtained with cameras located near the 
pilot's head, approximately 30 inches forward and 12 inches to the 
J 
4 NACA RM No. A8c25 
rear. The most distinct pictures were those taken with the camera 
directly in line with the shock wave . 
Background or Viewing Surface 
Some experimentation was carried on with background colors. 
Dark blue, light gray, red, and the polished aluminum surfaces all 
produced satisfactory results, although the best black and white 
photographs were obtained with the red surface. 
RESULTS AND DISCUSSION 
Location 
Shock waves have been observed on airplanes with both the 
conventional and low-drag airfoils, but photographs of the latter 
only have been obtained so far. Unpublished chordwise pressure-
distribution measurements made during tests of a similar airplane at 
the Langley Laboratory have been added to figure 4 to show that the 
shock-wave location indicated by the pressure distribution agrees 
~uite well with that photographed in flight. The pressure distribu-
tion was measured at the spanwise location of the plot in figure 4, 
which is somewhat outboard of the visible portion of the shock wave. 
Movement 
Below the Mach numbers at which severe buffeting occurred, the 
shock wave was observed to remain steady with Mach number and normal 
acceleration held constant. With increasing Mach number and constant 
acceleration, the shock-wave location progressed smoothly rearward, 
while, with increasing acceleration and constant Mach number, the 
location progressed forward. 
As the airplane was pushed to still higher Mach numbers and the 
pull-out begun, the shock wave would begin to oscillate fore and aft 
with an amplitude of 2 to 3 inches. During the oscillation, the 
general shift in location was observed to remain consistent with 
changes in Mach number and normal acceleration. In every case 
observed, the appearance of the oscillation was practically coincident 
with the incidence of severe buffeting of the airplane. As Mach 
number was decreased to the point where buffeting ceased, the shock 
wave would change position, moving forward steadily and decreasing 
in visibility until it disappeared when the Mach number fell below 
NACA RM No. A8C25 
the critical for the wing. Photographs illustrating this movement 
and diminishing visibility are presented in figure 5. The shock-
wave position in percent chor~which is also shown in this figure, 
was measured at a spanwise location of 28 percent of the semispan. 
5 
The oscillating shock wave was observed on two different 
airplanes and, in each cas~a definite relationship was felt by the 
pilot between the frequency of oscillation and of buffeting, although 
no quantitative measurements were obtained. 
Form and Variations Observed 
On the first test airplane, the wing shock-wave image took the 
form of a smooth curve with one or more reflexes, extending from 
about 18 inches from the wing root to within 24 inches of the wing 
tip. When oscillating, the entire shock wave appeared to move as a 
unit, although the outboard section would become relatively indistinct. 
This airplane had a smooth,painted surface and was exceptionally free 
from waviness in the wing skin. 
With a second airplane, having an unpainted wing and a rougher 
and more wavy surface, the shock wave app~ared broken into several 
sections rather than the smooth continuous curve mentioned previously. 
Oscillation of all visible segments took place simultaneously, 
although, again, the outboard position became indistinct. 
Canopy Shoyk Waves 
Photographs were also obtained of the image formed on the wing 
surface by refraction of the canopy shock waves, one of which is 
presented in figure 1. This is of particular interest because it 
produces the shock wave in profile, and clearly shows the form 
referred to as a lambda shock. Photographs were obtained which 
showed a definite fore and aft oscillation of this shock wave over 
the top of the canopy. 
ApplicatiOns 
Although these observations have been limited to three airplanes, 
it is felt that the method could be made applicable to any airplane, 
provided means of locating the camera overlooking the wing are avail-
able at a station almost directly opposite the shock wave. It is 
true, however, that, in airplanes where the pilot is located either 
6 NACA RM No. A8c25 
far forward or far aft of the wing station where the shock occurs, he 
will be unable to see the shock himself and so will encounter some 
di fficulty in placing the airplane correctly in relation to the sun. 
Practice in an airplane where he can visually observe the shock wave 
and the effect of the sun angle and airplane attitude on its visibility 
should make it possible for any pilot to set up the correct conditions, 
even though he is unable to observe the shock wave himself. 
Perhaps the greatest advantage of this method is its simplicity. 
Few instruments are re~uired, and very little practice should be 
re~uired of a competent pilot before he can attain good visible shock 
waves practically at will. 
The method should prove valuable in a further study of the 
relationship between airplane buffeting and shock-wave oscillation. 
A further extension of the method is indicated by the canopy shock 
photograph, which shows the shock wave in profile. By mounting the 
camera near the wing tip and flying with the sun directly opposite 
t hat wing, it should be possible to obtain profile shadowgraph images 
of the shock either against the side of the fuselage or against a low 
"fence" mounted at some inboard wing station. 
It is possible that future application may include obtaining 
photographs of the shock waves over other parts of the airplane 
(e.g., on wing-tip tanks as reported in reference 1). 
Possible Alternative Method 
In an attempt to obtain a visual indication of the wing shock 
position by another principle, a system of parallel dark lines was 
drawn over a white section of the wing. It was thought that viewing 
these lines through the wing shock wave WOuld, due to the differences 
in refraction between the air in and out of the shock wave, cause an 
apparent distortion of the lines at the wing-shock-wave intersection. 
Photographs obtained when the shock wave was known to be present did 
not reveal any noticeable distortion of these lines. Further develop-
ment, however, might show this method, which has the advantage of not 
re~uiring direct sunlight, to be usable. 
CONCLUSIONS 
The information contained in this report shows that visible 
shock waves may be observed over the wing and from the canopy of 
certain airplanes, using only the sun as a light source, if a given 
NACA RM No. A8C25 
set of conditions is strictly adhered to by the pilot. From the 
discussion and results, the following general conclusions may be 
drawn: 
1. The method will produce satisfactory shock-wave images 
t hat may be either viewed or photographed. 
2. Establishment of the correct test conditions in flight is 
relatively simple, re~uiring a minimum of practice by the pilot. 
7 
3. Extension of the method to include a study of the wing shock 
in profile, the shock over various other surfaces, and further study 
of the relationship between airplane buffeting and shock-wave oscilla-
tion all represent possible future applications. 
Ames Aeronautical Laboratory, 
National Advisory Committee for Aeronautics, 
Moffett Field, Calif. 
REFERENCE 
1. Johnson, Clarence L.: Development of the Lockheed P--80A Jet 
Fighter Airplane. Jour. Aero. Sci.,vol. 14, no. 12, Dec., 1947, 
pp. 659-679. 
-- -- ----------------------------------------

NAeA 1M No. A8C25 9 
Figure 1.- Examples of shock-wave images :photographed in flight. 

l 
NACA IM No. A8c25 
Low density 
medium 
~ Leading edge 
High density 
medium 
Upper surface 
of airfoil 
(a) Diagram of chordwise cross-section 
(b) Photograph equivalent to planform view of (a) 
11 
Figure 2 .- Illustration of method of viewing location of base 
of main compression shock wave on the wing of an air-
plane in flight utilizing the sun as the light source. 

NACA 1M No. A8c25 
Approximate permissible 
range of location of 
the sun--.... 
y 
13 
z 
x 
z 
Figure 3.- Proper orientation of the sun with respect to the 
airplane axes. 

NACA RM No. A8c25 
Chordwise pressure distribution 
from flight test results on a 
similar airplane. 
15 
-. /.o 
.5 
P 
oq 
Figure 4.- Comparison of shock-wave image position with pressure distribution 
measured in flight on a similar airplane. 

NACA 1M No. A8c25 17 
,. 
A -12498 
(a ) M = 0.74 . 
Figure 5.- Photographs of wing shock-wave movement in a dive recovery. 
1 
NACA EM No. A8C25 19 
(b) M = 0.69. 
Figure 5.- Conc1uQed. 
· I 
~~~ ___ J 


Visual observations of the shock wave in flight

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