A wake traversing probe was used to obtain section drag and wake profile data from the wing of a sailplane. The transducer sensed total pressure defect in the wake as well as freestream total pressure on both sides of the sensing element when the probe moved beyond the wake. Profiles of wake total pressure defects plotted as a function of distance above and below the trailing edge plane were averaged for calculating section drag coefficients for flights at low dynamic pressures

Saltzman, E. J.

English

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1 76 1101.0
5.3 Use of a Pitot Probe for Determlning
Wing Section Drag _n Flight
Edwin J. Saltzman
NASA Fllght Research Center
(A status report of development work of Lawrence C. Montoya and Paul F. Bikle)
Introduction
This paper presents the status of recently completed development work on the
wake traverse method of obtaining section drag at low speeds. The method of B. M.
,Jones, Reference 1, has been applied to wake profile data obtained from the wing of
a sailplane.
Though the sailplane provides a quiet and relatively vibration free environment,
and a very clean and smooth high aspect ratio wing, it is limited to a dynamic pressure
range from about 6 psf to somewhat above 20 psf. These low dynamic pressures are a
severe challenge for the wake traverse method.
This paper is intended not only to rekindle interest in the wake traverse approach
of defining profile drag, but it is also intended to demonstrate techniques for increasing
reliability and minimizing certain bias errors when dealing with relatively low differen-
tial pressures. The thought which accompanies this paper is that if acceptable accuracy
in section drag can be obtained at these low dynamic pressures, even better profile drag
definition could perhaps be obtained for applications on general aviation aircraft where
higher pressures would prevail.
Airplane and Test Cond|t;ons
The sailplane was a T-6 having a modified Wortmann FX-61-163 airfoil. The
modification consisted of a straightening of the underside cusp region at the rear portion
of the section. A sketch of the wing profile is shown in Figure 1.
The surface finish of the wing was very smooth and the maximum waviness was
about 0.003 inches in a two-lnch section of surface. At the semlspan station of the
wake traverse tests the waviness was less than this.
The wake measurements were made 9.6 inches behind the wing trailing edge
which corresponds to about 32 percent of the 29.9 inch local chord. Data were
obtained for speeds from about 40 knots to 125 knots which Provided chord Reynolds
numbers between 106 and 3 x 106.
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Indicated airspeed was held constant during each data run and six "total-pressure"
wake traverses were made during each run followed by three "statlc-pressure" wake
traverses. This sequence of nine traverses took approximately one minute. Data
been obtained for deflected flap conditions but the data to be shown herein are for zero
flap deflection. The majority of the data were obtained for very smooth air conditions
but some data were obtained for air that would be considered somewhat rough.
A photograph of the sailplane with the wake traverse probe installed is shown
in Figure 2. A Kiel probe can be seen mounted ahead of the drive unit package on the
right wing. The dark colored tape running parallel to the wing trailing edge covers and
holds down the wlrlng harness which connects the drive unit package to the recording
package inside the fuselage.
Instrumentation
Probe and Drive Unit: A closer view of the probe and drive unit, and some
of the reference probes, is provlded in Figure 3. The Kiel tube is used as the reference
for the transversing total pressure probe and the trailing boom provides a static reference
for the traversing statlc pressure probe. The trailing boom orifices are located about 5
feet behind the wing and are calibrated against the ships static system, which is in turn
a system whlch has been calibrated for posltion error.
The traversing probe (with total and static heads) and drive unit are shown
again in Figure 4(a) and in closer detail in Figure 4(b) where some of the parts have
been labeled. The probe traversed to about 8 inches above and below the wing trailing
edge. The probe travel rate was a little less than 3 inches per second. The hardware
which is shown in Figure 4 weighed about 3 pounds.
A very important part of the unit was the switching valve wh ich permitted the
same pressure transducer to measure the wake station total pressure decrement, in one
mode, and the difference between wake static and traillng boom static pressure, when
in another mode. More detail will be provided about this feature in following figures.
Recorder Package: The recording package consisted of a tape recorder, battery
and a component box which housed a pressure transducer and two amplifiers. This
package was mounted on a shelf behind the pilot's head rest as can be seen in Figure 5.
A closer view of this hardware is shown in Figure 6, after a covering hatch has been
removed. The weight of this package is about 40 pounds. The switch shown in Figure 6
is also an important element in obtaining in-flight tare readings on the transducer
which records "ships q."
172
Switching Features: Figure 7 will be used to show schematically how the
switching feature mentioned in the previous section works. As sketched in Figure 7,
the switch is in the "zeroing mode." When the pilot places the switch in this mode,
freestream total pressure exists on each side of the pressure transducer element, thus
providing an in-fllght tare reading which minimizes the bias error of this transducer.
When the switch is placed in the "q mode" (toggle to the left) there is fi'eestream total
pressure on one side of the transducer and ships static on the other side. This, of
course, provides a record of "ships q." Because the ships static system has been cali-
brated for position error the appropriate corrections are applied and a true freestream
dynamic pressure can be calculated as a function of time for correlation with the
pressures recorded in the wake and the probe position in the wake. The airspeed
indicator allows the pilot to hold "ships q" steady for a sufficient period of time to
permit up to nine successive wake traverses to be made under quasl-steady state
conditions.
Another switching feature involves the switching valve which was identified
in Figure 4(b). In Figure 8t a slide valve is shown in schematic form to illustrate how
this switching valve can be used to direct Kiel tube pressure (freestream total) to one
side of the transducer and wake total pressure to the other side. Thus the transducer
senses the total pressure defect, AP T, in the wake. When the wake probe moves
beyond the wake, the transducer experiences freestream total pressure on both sides
of the sensing element which thereby provides inflight tare readings for the transducer.
This feature minimizes the bias error for this transducer.
Another made for this switching valve is illustrated in Figure 9. In this case
the valve arrangement provides trailing boom static pressure to one side of the sensing
element and wake station static pressure on the other side. As mentioned before, the
trailing boom static pressure has been calibrated against the ships static system which
in turn has been calibrated for position error. Therefore the true decrement between
wake station static pressure and freestream static can be calculated. It is this corrected
decrement, AP, which appears as an adjustment in the Jones expression for calculating
section drag from wake measurements. The Jones equation follows:
- 2 S _J qv_ake ( ,_l/ wake "l" "kl<P)Cd - _- 1- q ' dyq
where
L
q _ freestream dynamic pressure, from ships transducer plus position error
correction
173
A PTwak e , total pressure decrement in wake, from probe transducer
AP , difference between freestream and wake static pressure, from probe
transducer
dy, from probe position potentlometer
c, wing chord at wake survey location
qwake + AP , fromq -APT
wake
qwake , from (qwake + _ P ) - A P
or (q - APTwak e) - AP
It is important that the same transducer, through the switching valve feature lust
described, provides both _PTwake, and AP because thereby the parameter AP
has bias error minimization in the same way as has been described for the A PTwak e
measurement.
It is also important to note that six APTwak e traverses and three AP traverses
are made in succession while indicated airspeed, and consequently q , ;s held constant.
Section drag coefflcl ents to be presented in a following section are the average of six
such traverses for each Reynolds number condition.
Resul ts
Profiles of APTwak e plotted as a function of distance above and below the
wing trailing edge plane are presented in Figure 10. These are obtained from six
consecutive traverses through the wake over a period of about 40 seconds. The airspeed
was about 44 knots which resulted in a section lift coefficient of about 1.0. Note the
low "delta pressures" with which the instrumentation must contend at these speeds.
Some part of the apparent dispersion of the six profiles is random scatter; however,
a part of it is caused by small changes in airspeed during the 40 seconds of tun time.
Therefore, a part of the apparent dispersion will be eliminated when each profile is
normalized by the appropriate mean q for that traverse. As mentioned before, the six
normalized profiles are then averaged when calculating section drag coefficients by
,Jones' equatlon_.
Another set of profiles is shown in Figure 11 for an airspeed of about 42 knots
which provided a llft coefficient near maximum. In this case only, four profiles are
174
shown. Theprofile with triangular symbolsis the first of the sequence and is similar
in appearance to those shown in Figure 10. The following profiles in their order of
occurrence (diamonds and squares) progress toward local stall (circles) as speed was
reduced ever so slightly. Earlier flying with flow visualization indicated that this
region of the wing was indeed stalling at these speeds.
Section drag coefficients plotted as a function of chord Reynolds number are
shown in Figure 12. Also shown are curves from Blasius and Schlichting representing
fully laminar and turbulent flow over the upper and lower surfaces as if they were flat
plates.
The circles represent three flights (each point is the average of 6 traverses) for
natural transition. The wing surface was exceptionally smooth and clean for these runs.
The squares represent flight with the boundary layer tripped 5 percent behind
the leading edge for both the upper and lower surfaces. The trip material was dis-
tributedgrit of 0.035 inches mean height above the skin surface.
The results obtained to date have been encouraging and suggest several
possibilities which may deserve to be included in follow-on flights. Some of these
possibilities are listed below:
(a) 400 grit sanding of the local surfaces
(b) 200 grit sanding of the local surfaces
(c) add waves, up to 0.008" per 2 inches
(d) restore to best finish (fill, sand, rub)
(e) tape over flap gap
(f) _nsect roughness (bugs) on leading edges
(g) install wake traverse unit on several general aviation aircraft .
Closing Remarks
The experience reported has provided progressively improving accuracy in defining
section drag for flight at low dynamic pressures. The accuracy is about 3 to 4 percent for
the lower dynamic pressures and somewhat better than 3 percent at the higher dynamic
pressures. It should be emphasized, however, that such accuracies were obtainable only
by resorting to the detailed procedures described herein (especially the in-flight switching
and tare evaluations) and by conducting all phases of the testing with the utmost care.
Also of great importance was the precise position error calibration which had been
accomplished on the ships airspeed-altitude system prior to the wake traverse work.
175
It is believed that with proper care and attention to detail, comparable or
perhaps even more accurate section drag coefficients could be obtained on general
aviation aircraft, considering the higher dynamic pressures encountered in general
aviation.
References
Reference 1. Jones, B. M., Measurement of Profile Drag by the Pitot-
Traverse Method, Cambridge University Aeronautics Laboratory, Rand M.
No. 1688, 1936.
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Use of a pitot probe for determining wing section drag in flight

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