A limited calibration of a combined pitot-static tube and vane-type flow-angularity indicator has been made in the Langley 4- by 4-foot supersonic pressure tunnel at Mach numbers of 1.61 and 2.01. The results indicated that the angle-of-yaw indications were affected by unsymmetric shock effects at low angles of attack

Mace, William D

Sinclair, Archibald R

English

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NATIONAL ADVISORY COMMITTEE 
FOR AERONAUTICS 
TECHNICAL NOTE 3808

WIND-TUNNEL CALIBRATION OF A COMBINED PITOT-STATIC TUBE 
AND VANE-TYPE FLOW-ANGULARITY INDICATOR AT MACH

NUMBERS OF 1.61 AND 2.01 
By Archibald R. Sinclair and William D. Mace 
Langley Aeronautical Laboratory
Langley Field, Va.
77 
Washington 
October 1956
https://ntrs.nasa.gov/search.jsp?R=19930084583 2020-06-17T18:36:16+00:00Z
NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
TECH['TICAL NOTE 3808 
WIND-TUNNEL CALIBRATION OF A COMBINED PITOT-STATIC TUBE 
AND VANE-TYPE FLOW-ANGULARITY INDICATOR AT MACH 
NUvIBERS OF 1.61 AND 2.01 
By Archibald R. Sinclair and William D. Mace 
SUVIMARY 
A limited calibration of a combined pitot-static tube and vane-type 
flow-angularity indicator has been made in the Langley 4-'by 4-foot super-
sonic pressure tunnel at Mach numbers of 1.61 and 2.01. The results indi-
cate that the instrument registers too high an angle of attack and gives 
an error of 0.70 at an angle of attack of 200 for a Mach number of 1.61 
and an error of 1.60 at an angle of attack of 240 for a Mach number of 
2.01. At zero angle of attack the flow field about the yaw vane was unsym-
metrical and caused an error of i.i-° in yaw indication at zero angle of 
yaw for a Mach number of 2.01. The installation of a dummy vane pedestal 
to provide a more symmetrical flow field reduced this error to 0.250. The 
probe gave static-pressure readings which were too low at angles of yaw. 
INTRODUCTION 
A combined pitot-static tube and vane-type flow-angularity indicator 
has been designed by the National Advisory Committee for Aeronautics for 
use on research aircraft. This instrument is designed to measure impact 
and static pressures by means of a pitot-static tube and to measure angles 
of attack and yaw by means of two free-floating vanes. The purpose of 
the present investigation in the Langley Ls. by 4-foot supersonic pressure 
tunnel was primarily to obtain an aerodynamic calibration of the angle-
of-attack vane at supersonic speeds, since this calibration was required 
in order to evaluate the drag characteristics of supersonic research air-
craft. Tests were made at Mach numbers of 1.61 and 2.01 and at pressure 
altitudes of 30,000 feet and 60,000 feet. A limited amount of calibra-
tion data were also obtained for the angle-of-yaw vane. The effect of 
yaw on the indicated static pressure is also presented. The effective-
ness of a dummy vane pedestal in producing a more symmetrical flow field 
about the angle-of-yaw vane was also investigated.
2
	
NACA TN 3808
SYMBOLS 
M	 stream Mach number 
p	 free-stream static pressure 
Pi	 free-stream static pressure as indicated by instrument 
a	 stream total pressure minus stream static pressure 
a	 angle of attack, deg 
ii	 angle of yaw, deg
APPARATUS AND TESTS 
The instrument which was calibrated in this investigation (fig. i) 
consisted of a pitot-static tube for indicating impact and static pres-
sures and free-swiveling vanes for indicating angles of attack and yaw. 
This instrument is similar to the one previously calibrated in the Langley 
8-foot transonic tunnel. (See ref. 1.) The major differences are in the 
shape of the pitot tube and in the dimensions of the vanes. In the pres-
ent instrument the vanes have no stops and can rotate a full 3600. 
Details of the pitot-static component of the instrument are shown 
in figure 2. There are two identical groups of static orifices, forward 
and rearward, and each group is brought out independently. There are a 
larger number of orifices on the bottom of the tube than on the top in 
order to reduce the static-pressure error at positive angles of attack. 
The angularity-indicating component consists of free-swiveling mass-
balanced vanes located downstream of the pitot-static component on the 
enlarged part of the instrument, as shown in figure 1. For some tests, 
a dummy vane pedestal was mounted opposite the angle-of-attack vane ped-
estal as shown by the dashed outline. 
When this system is used in flight, vane position is measured by 
means of a syncbro-transmitter system connected to the vanes. For the 
calibration in the wind tunnel, this system could not be used directly 
because of variable and unknown deflections of the probe and the support 
system under aerodynamic loading. In order to eliminate the effect of 
these deflections, an optical system used in conjunction with a small 
mirror mounted flush in one of the vanes was used to measure the angle 
of the vane relative to the tunnel center line. The vane error could
NACA TN 3808
	
3 
thus be measured directly, independently of any deflections, and with a 
high degree of accuracy. For the calibration of angle of yaw with varia-
tion of angle of attack at M = 2.01, the data from the synchro transmitter 
were faired through data points obtained optically. The maximum error in 
the optically measured angles is believed to be less than 0.10. 
For the tunnel tests, the instrument was mounted on an adapter as 
shown in figure 3 . This adapter, in conjunction with the regular model 
support sting, was designed to maintain the angle-of-attack vane axis at 
very nearly the same point near the tunnel center line when the probe 
angle was changed in the horizontal plane. The adapter-sting combina-
tion was fitted with several joints permitting a 3600 rotation of the 
probe; thereby, either the angle-of-attack vane or the yaw vane could be 
placed in a horizontal plane containing the tunnel axis. 
Tests were made at stream static pressures corresponding to alti-
tudes of 30,000 feet and 60,000 feet at a Mach number of 1.61 and to 
an altitude of 60,000 feet at a Mach number of 2.01. 
At M = 1.61, tests were made with the probe fixed at a = 0 to 
check the repeatability of different vanes. For this purpose, tests 
were made with two different vanes chosen at random and with the vane 
in which the mirror was mounted. It was necessary to use the servo-
transmitter system to read the vane position; but, since the probe was 
not moved, it can be assumed that the probe deflection was the same in 
all three tests. No significant differences were found in the trail 
angles of the three different vanes. 
For both the angle-of-attack and the angle-of-yaw tests, repeat 
tests were made with the probe rotated 180 0 to eliminate the effect of 
any tunnel flow angularity. 
At each Mach number, a transverse survey of the test section was 
made by moving the probe across the test section at a fixed angle of 
attack of 50 while the angular position of the mirror-fitted vane was 
measured optically. The variation in flow angularity was found to be 
p1 
- p0 less than 0.10. The maximum variation in
	
	 was 0.003 at M = 1.61 
qc 
and 0.001 at M = 2.01. 
Probe static pressures were measured, but for the angle-of-attack 
tests the results showed only random scatter with no significant trends. 
The explanation of this scatter is not entirely known, but it is probably 
partly a result of the pitot-static part of the probe moving in a large 
arc across the forward part of the test section where there may have been 
flow variations in regions not covered by the transverse survey. These 
angle-of-attack pressure data are, consequently, not presented. Varia-
tion of probe static pressure with yaw, however, was of sufficient mag-
nitude to show a systematic trend and is presented for M = 2.0.
1.	 NACA TN 3808 
RESULTS AND DISCUSSION 
The results of the angle-of-attack correction at zero angle of yaw 
are shown in figure 1, where the angle-of-attack correction (the true 
value of a. minus the measured value of a.) is plotted against angle of 
attack. These results show that at zero angle of attack the probe indi-
cates correctly but, as a. increases, the probe indicates too high an 
angle. This error reaches a maximum of about 0.70 at a. = 200 for 
M = 1.61 and 1.60 at a. = 240 for M 2.01. This error is probably 
a result of upwash around the inclined cylindrical part of the probe. 
The effect of altitude is within the accuracy of the measurements except 
possibly at a. = 200, where the error is about 0.1 0 less at 60,000 feet 
than it is at 30,000 feet for M = 1.61. 
The variation of angle-of-yaw correction (the true value of it minus 
the measured value of ) with angle of yaw for M = 2.01 is shown in 
figure 5. At zero angle of attack (fig. 5(a)) the correction at zero 
angle of yaw was li .°. This large error probably results from the unsym-
metrical effects of the strong detached shock ahead of the angle-of-attack 
vane pedestal. In an effort to make the flow field about the angle-of-
yaw vane more symmetrical, a dummy pedestal (see fig. 1) was installed 
opposite the angle-of-attack vane pedestal. As shown by the dashed curve 
in figure 5(a), the resulting error curve is more nearly symmetrical. 
The correction at zero angle of. yaw was reduced from 1.4 0 to 0.250 . Sim-
ilar yaw data for an angle of attack of 9.8° are shown in figure 5(b). 
For this angle of attack the yaw vane was apparently out of the influence 
of the unsymmetrical flow field, since the yaw correction Is essentially 
zero both with and without the dummy pedestal. Only limited optical data 
could be obtained with this configuration because of the geometry of the 
installation. 
The variation of angle-of-yaw correction with angle of attack at 
zero angle of yaw and a Mach number of 2.01 is shown in figure 6. Optical 
measurements were obtained for only two angles of attack, 00 and 9.80, as 
indicated by the symbols. The curves through these points were obtained 
by adjusting the data obtained from the synchro-transmitter system so 
that the curves passed through the points obtained by optical measure-
ment. The method used in adjusting these data was based on the assumption 
that the errors in the synchro-transmitter data were a result of an angu-
lar deflection of the probe; probe deflection was assumed to be a linear 
function of angle of attack. Some error probably results from this 
assumption, but the trends shown are probably correct. From these results 
it appears that the unsymmetrical-flow effects on the angle-of-yaw vane 
largely disappear at angles of attack above about 7. 
The variation of measured probe static pressure with angle of yaw 
is shown in figure 7. The probe static pressure showed a rapid drop 
with an Increase in angle of yaw, as would be expected with a probe
NACA TN 3808
	
5 
having orifices on only the top and the bottom of the tube. The indicated 
static pressure is about 5 percent too low in terms of q at 4i = 200. 
CONCLUSIONS 
A wind-tunnel calibration of a combined pitot-static tube and vane-
type flow-angularity indicator at Mach numbers of 1.61 and 2.01 has led 
to the following conclusions: 
1. At zero angle of yaw, the instrument registers too high an angle 
of attack, except at an angle of attack of zero. The error was a func-
tion of Mach number and increased with angle of attack to a maximum of 
about 0.70 at an angle of attack of 20 0
 for a Mach number of 1.61 and 
1.60 at an angle of attack of 240
 for a Mach number of 2.01. 
2. The unsymmetrical flow, field from the angle-of-attack vane pedestal 
caused large errors in the yaw-vane indications at an angle of attack of 
zero. This error reached a maximum of 1i40
 at zero angle of yaw at a 
Mach number of 2.01. Installation of a dummy pedestal to give a more 
symmetrical flow field reduced this error to 0.25 0 . The unsymmetrical 
flow effects largely disappeared at angles of attack greater than about 7. 
3. The static-pressure tube with orifices on the top and bottom gave 
readings which were too low at angles of yaw. 
Langley Aeronautical Laboratory, 
National Advisory Committee for Aeronautics,

Langley Field, Va., June 14, 1956. 
1. Pearson, Albin 0., and Brown, Harold A.: Calibration of a Combined 
Pitot-Static Tube and Vane-Type Flow Angularity Indicator at Tran-
sonic Speeds and at Large Angles of Attack or Yaw. NACA 14 L52F24, 
1952.
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= 00.
10
	 NACA TN 38o8 
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NACA TN 3808
	
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Figure 6.-.. Variation of angle-of-yaw correction with angle of attack.

M = 2.01; pressure altitude, 60,0oo feet; 4r = 00. 
4', deg 
Figure 7 . - Variation of measured probe static pressure with angle of yaw. 
M = 2.01; pressure altitude, 60,000 feet; a. = 00 Pressures obtained 
from forward orifices. 
NACA - Langley Field, Va.


Wind-tunnel calibration of a combined pitot-static tube and vane-type flow-angularity indicator at Mach numbers of 1.61 and 2.01

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