The suggestion for the measurements collected in this report was prompted by the results of spinning experiments. For the theoretical spinning investigations there was a lack of data on lateral force, yawing, and rolling moment of airplanes

Blenk, Hermann

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TCHNI CAL 
ITAT IONAL ADVISORY COLCvIITTEE FOR ARCNAVTI OS 
No. 586 
GOTTINEN sIx-COLPCI•TE::T SCALE IASUPLS ON 
A JTJTKRS A 35 AIRFLAE hODEL 
By Herniann Elenk 
Jahrbuch 1930
der Deutschen Versuchsanstalt ±'r Luftfahrt 
Wash ingt on
October, 1930
FILE COpy 
ToLmr8turflto 
the fflez of tim Natj.naj 
Mvisory Gomrrwtt 
for A8ronaut 
Wasiingto D. C.
https://ntrs.nasa.gov/search.jsp?R=19930094830 2020-06-17T00:25:56+00:00Z
NATIONAL ADVISORY COivflITTEE FOR AERONAUTICS. 
TECHNICAL MEMORANDUM NO. 586. 
GTTINGEN SIX-COMPONETT SCALE MEASUREMENTS ON 
A JUNKERS A 35 AIRPLANE MODEL. 
By Hermann Blenk. 
I. Introduction 
This is a report on the measurements made by the Ggttingen 
Aerodynamic Laboratory, at the request of the D.V.L. (Deutschen 
Versichsanstalt fr Luftfahrt). The data are given here in 
condensed form. 
The suggestion for these measurements, as for the previ-
ously reported measurements on dihedral, sweepback, and warping 
(Luftfahrtforschung, Vol. III, 1929, p. 27), was prompted by 
the results of spinning experiments. For the theoretical spin-
ning investigations there was a lack of data on lateral force, 
yawing, and rolling moment of airplanes (aside from English 
and American experiments). In order to be able to compare the 
model tests themselves and the theoretical investigations based 
upon theni with the flight practice, we chose a niodel of the 
well-known Junkers A 35 - low-wing type monoplane (Fig. 1). 
The scale of the model was 0.085, without propeller. Further 
tests with running propeller are to be published later. 
* Il Gottinger Sechskomponentenmessungen an einem Mod.ell des Flug-
zeugmusters Junkers A 35." From 1930 Yearbook. of , the D.V.L.
'I 
N.A.C.A. Technical Memorandum No. 586
	
2 
II. Definitions 
The following symbols are used: 
.	 geometrical angle of attack, 
Y	 angle of Yaw, 
angle of aileron deflection, 
H angle of elevator deflection, 
s angle of rudder deflection, 
F wing area, 
b wing span, 
t	 wing chord (in center), 
q	 static pressure, 
A	 lift, its coefficient c = 
a qF 
W	 drag, its coefficient cw = 
S	 cross force, its coefficient cs = 
1'h pitching moment, its coefficient Cmh = 
1	 yawing moment, its coefficient cms = 
Mq rolling moment, its coefficient cmq 
Figure 2 shows the positive angles, forces and moments. 
The x-axis of the coordinate system forms the usual reference 
line of the wing. The angle of the air f]ow defines the x-axis 
so that the wind direction falls in the x-y plane. The angle
NIA.C.A. Technical Memorandurri No. 586 	 3 
formed by wind direction and x-axis is the angle of yaw y. 
The z-axis is perpendicular to the other two axes. Axes x, y, 
and z form a clockwise system. The angle of the mean wing 
chord and a line parallel to the y-axis is the angle of attack a 
Lift A, drag W, and cross-wind force S are positive 
in the direction of positive x, y, and z axes. The moments Mh, 
M8 , Sand Mq, are positive when turning negatively in the direc-. 
tion of the clockwise system of axes. The positive directions 
of rotation of control movements H,	 and	 have been 
chosen so that positive control movements produce positive mo-
ment S.
III. Synopsis of Measurements 
We made three-component scale measurements: 
1) On wing alone for angle of attack ranging from a = _200 
to +900; angle of yaw 'Y = o°; and ailron angle	 = 00. 
2) On whole airplane for a = -.2O 0 to +90°; angle of yaw 
= 0 ; and aileron angle 6ç =	
= s = 00. 
3) On whole airplane for a = -6° to +15°; angle of yaw 
= 00; aileron angle
=
= 0 0 ; ± 100.
4) On fuselage with tail surfaces minus landing gear, for 
a = - 10° to +450; angle of yaw y = 0°; rudder angle
	
= 0°; 
= 0°; and ± 10°. 
N.A.C.AS Technical Memorandum No. 586	 4 
Six-component scale measurements were made on the whole 
airplane for	 =	 70 to + 70, and for each four and two an-
gles ofattack, respectively; the combinations of the control 
angles were: 
s	
00	 ç= 00 
-5° 00 5° 
00 15° 
_50 50 0° 
, 50 50 50 
_50 50 15° 
-5° 100 00 
-5° 100 50 
- 5° 10° 15° 
IV. Results of Tests
Figures 3 and 4 show the polars and the pitching moment 
curve for the A 35 wing at a. = - 200 to + 90 0
 and 'Y = 00. 
The curves corresponding to the whole airplane are shown in Fig-
ures 4 and 5 with
	
dy.,:	
H' and	 s = 00 . It should be 
noted that the curves in Figures 3 and 5 are made at a different 
scale from those of Figure 4. Figure 5 also contains the polars 
and. the pitching moment curves of the whole airplane with ele-
vator settings of -10° and +10°, while Figure 6 gives the data 
for the fuselage with tail surfaces without landing gear. The 
C a, CW and °mh coefficients here are plotted against the wing 
N.A.C.AU Technical Memorandum No. 586 	 5 
area and wing chord, while the angle of attack is plotted against 
the propeller axis. The incidence of the mean chord to the pro-
peller axis is 3°. The longitudinal moment (for fuselage with 
tail, Fig. 6) is measured about the lateral axis through the 
intersection of thrust axis and nose of fuselage. By comparing 
the respective measurements on the whole airplane, on the wing 
alone, and on the fuselage with tail surfaces, we can draw con-
clusions as to the amount of downwash at the tail surface. 
For known reasons the six-component scale measurements 
(Figs. 7-15) had to be confined to angles of yaw 	 = - 70 to 
+ 70• It was found that lift, drag and longitudinal (pitching) 
moment (values from usual three-component scale measurements) 
are independent within the measured range of ',	 , and	
. 
For this reason we can omit in the following, the data on ca, 
Cw and c; they can, moreover, be quite accurately determined 
from Figures 4 and 5. This enables us to present the six-compo-
nent scale measurements more simplified. 
In six-component scale measurement., the points of measure-
ment are spread out considerably, so in order to make it as ex-
pressive as p ossible, we omitted the compensation curves and. 
merely connected the points of measurement by straight lines. 
Figures 7, 8, and 9 show the coefficients of the cross-wind 
force c	 of the whole airplane plotted against the angle of 
yaw 'y (Fig. 7: H = - 5°, and	 = 00; Fig. 8:	 = 5°; 
Fig. 9:	 = 150 ). As y' increases, S decreases. Hence the
N.A.C.A. Technical Memorandum No. 586 	 6 
cs 
value	 is negative and practically unaffected by 	 , and 
Even the effect ot angle a on is generally small. 
A rudder movemen; pduces a positive cross-wind force, whose 
value, within the measured range of all other angles (a, y, 
is practically unaffected. This is seen in the graphs, 
where the corresponding curves for different cross-wind angles 
run parallel at the same distance (measured in the ordinate axis). 
Comparison with the cross-wind forces measured on the wing alone, 
which, within the range in question, amounts to c 5 = 0.01, show 
that by far the greater part of the cross-wind force was con-
tributed by tail surfaces and fuselage, as we expected. 
Figures 10 to 12 give the coefficients of yawing moment 
cms plotted against y (Fig. 10:	 = -°, and	 = 0 0 ; Fig. 
11:	 = 50; Fig. 12:	 = 15°). The moment decreases as Y 
increases. The value 	 cms/V likewise is negative and prac-
tically unaffected by 	 and	 On the other hand, the ef-
fect of a on	 crns/	 is at times quite enormous, which agrees 
with the previously mentioned measurements on individual wings. 
A rudder movement produces a positive lateral moment, but an 
aileron movement produces a negative lateral moment. Both con-
trols work against each other in this respect. At a = 14.50, 
for example, the effect of a 	 = 10° rudder deflection on the 
lateral moment is practically counterbalanced by .a	 = 50 
aileron movement. The effect of an aileron movement on the yaw-
ing moment increases with the angle of attack, in conformity
N.A.C.A. Technical Memorandum No. 586	 7 
with the theory of Betz and Petersohn.* A numerical comparison 
with the repeatedly cited measurements on individual wings shows 
the order of magnitude of the yawing moment to be about the same 
in both cases. So we conclude that a large portion of the yaw-
ing moment is contributed by the wings (in contrast with the 
cOss-wind force). 
Figures 13 to 15 contain the coefficients o' the rolling 
moment of the whole airplane plotted against angle 'y (Fig. 13: 
= _5P ,
 and	 = 0 0 ; Fig. 14: s = 5°; Fig. 15: s = 100). 
While •the measurements for constant aileron deflection 	 were
combined in Figures 7 to 12, each diagra here shows the meas-
urements for constant	 . The rolling moment increases as ' 
increases. Et the increase in Cmq/Y is almost negligible 
at small or average a, and increases to a multiple only when 
approaching the burbling point, a phenomenon very important for 
stalling flight. An aileron deflection 	 produces a strong
poitive rolling moment. In our example, rolling moments pro-
duced by a
	
= 5° aileron deflection, are larger than those 
occurring at a = 14.50 , ' = 7°, and	 = 00 , so that within 
* A. Betz 1 and E. Petersohn, "Zur Theorie der Querruder, " Zeit-
schrift fur angewandte Mathematik und Mechanik, Aug. , 1928, Vol. 
VIII, pp. 253-257. (See N.A.C.A. Technical Memorandum No. 542: 
"Contribution to the Aileron Theory," Dec., 1929.) 
C. Wieseleberger, "Theoretische Untersuchungen ber die Querru-
derwirkung beim Tragflugel." Report of the Aeronautical Research 
Institute, Tokio Imperial Univrsity, Dec., 1927, No. 30, Vol. II, 
No. 16, pp. 421-446. (See N.A.C.A. Technical Memo r andum iTo. 610: 
"Theoretical investigation of the Effect of the Ailerons on the 
Wing of an Airplane," April, 1929.)
N.A.CSA. Technical Memorandum No. 686
	
8 
this range the control of rolling seems to be assured with the 
aileron. With stronger side wind, however, the conditions be-
come less favorable. The deflection of rudder	 has no ef-

fect on the rolling moment; Figures 13 to 15 almost correspond, 
as we expected. From this and from a comparison with the wing 
measurements we can conclude that the rolling moment is pro-
duced by the wing alone. 
Translation by J. Vanier, 
National Advisory Committee 
for Aeronautics.
15.94 in I 
3.50 m 
r = 0.8 rn
3.50 m 
N.A.C.A. Technical Memorandum No.586	 Fig.l 
C\2 
c':2
t=2.20m	 t1.SOm 
F = 29.76 rn2
	
t 
T1 - _____ - 
G = 1600 kg 
- 
— b	 15.94 i--------i
Fig.l The Junkers A 35 airplane.
y 
S 
Wind 
x	
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Fig.2 
N. A .0 .A. Technical Memoraidum No.586	 Figs.2, 3 
0	 -0.2	 -0.4	 -0.6 
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h,	 2.7 
1,	 5.6 
j, 8.6 
k, 11.5 
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rn, 14.5 
n, 17.b 
a, 20.5 
p, 20.5 
q, 23.0 
r, 27.6 
s, 32.70 
t, 35.2 
U, 37.7
0 
v, 42.7 
w, 47.7 
x, 54.7 
y, 59.8 
z, 64.8 
a 89.8 
b 74.9 
79.9° 
d 84.9 
e 90.0° 
Fig. 3 
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cxJNcOC\2 
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N.A.C.A. Techiical Memorandum No.586	 Fig.4 
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(UNNaDO)	 - 
N.A.C.A Technical Memorandum No.586
	
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Gottingen six-component scale measurements on a Junkers A 35 airplane model

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