Flight investigation at high mach numbers of several methods of measuring static pressure on an airplane wing

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Material Information

Title:
Flight investigation at high mach numbers of several methods of measuring static pressure on an airplane wing
Alternate Title:
NACA wartime reports
Physical Description:
9, 9 p. : ill. ; 28 cm.
Language:
English
Creator:
Zalovcik, John A
Daum, Fred L
Langley Aeronautical Laboratory
United States -- National Advisory Committee for Aeronautics
Publisher:
Langley Memorial Aeronautical Laboratory
Place of Publication:
Langley Field, VA
Publication Date:

Subjects

Subjects / Keywords:
Mach number   ( lcsh )
Bombers   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: A flight investigation was made to compare static pressures in subsonic and supersonic flow over an airplane wing as measured by static-pressure tubes, a static-pressure belt, and orifices flush with the wing surface. The measurements were made on the upper surface of the wing of the P-47D airplane over a range of flight conditions in which local Mach numbers from 0.34 to 1.41 were obtained at the measurement stations. For some of the tests, a total-pressure tube was mounted on the wing surface to determine its characteristics in supersonic flow. The results indicated that static-pressure measurements obtained with suitably designed and installed flush orifices, static-pressure tubes, and static-pressure belt will be in reasonable agreement for both subsonic and supersonic flow. The pressures in supersonic flow measured by the total-pressure tube mounted on the wing surface were found to be in close agreement with values predicted by theory.
Bibliography:
Includes bibliographic references (p. 9).
Statement of Responsibility:
by John A. Zalovcik and Fred L. Daum.
General Note:
"Report no. L-90."
General Note:
"Originally issued November 1944 as Restricted Bulletin L4H10a."
General Note:
"Report date November 1944."
General Note:
"NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance research results to an authorized group requiring them for the war effort. They were previously held under a security status but are now unclassified. Some of these reports were not technically edited. All have been reproduced without change in order to expedite general distribution."

Record Information

Source Institution:
University of Florida
Rights Management:
All applicable rights reserved by the source institution and holding location.
Resource Identifier:
aleph - 003613532
oclc - 71212988
sobekcm - AA00006301_00001
System ID:
AA00006301:00001

Full Text


RB No. L4H10Oa


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS






WAR1lTIME RE PORT
ORIGINALLY ISSUED
November 1944 as
Restricted Bulletin L4HLOa

FLIGHT INVESTIGATION AT HIGH MACH NUMBERS

OF SEVERAL MEIHOIS OF MEASURING STATIC
PRESSURE ON AN AIRPLANE WIMG

By John A. Zalovcik and Fred L. Daum

Langley Memorial Aeronautical Laboratory
Langley Field, Va.


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WASHINGTON

SNACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of
advance research results to an authorized group requiring them for the war effort. They were pre-
Sviously held under a security status but are now unclassified. Some of these reports were not tech-
nically edited. All have been reproduced without change in order to expedite general distribution.


L 90


DOCUMENTS DEPARTMENT






































Digitized by the Internet Aichive
in 2011 with funding from
University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation


http://www.archive.org/details/flightinvestigalang









NACA RB Yo. L41Oa


NATIONAL ADVISORY COM7ITTTEE FOR AERONAUTICS


RESTRICTED BULLETIN


FLIGHT INVESTIGATION AT HIGH MACH NUMBERS

OF SEVERAL METHODS OF 'rEASURI!TG STATIC

PRESSURE ON AN AIRPLANE 'TJING

By John A. Zalovcik and Fred L. Daum


SUIsTARY


A flight investigation was made to compare static
pressures in subsonic and supersonic flow over an air-
plane wing as measured by static-pressure tubes, a
static-pressure belt, and orifices flush with the wing
surface. The measurements were made on the upper surface
of the wing of the -L7D airplane over a range of flight
conditions in which local !'ach numbers from 0.54 to 1.41
were obtained at the measurement stations. For some of
the tests, a total-pressure tube was mounted on the wing
surface to determine its characteristics in supersonic
flow.

The results indicated that static-pressure measure-
ments obtained with suitably designed and installed flush
orifices, static-pressure tubes, and static-pressure belt
will be in reasonable agreement for both subsonic and
supersonic flow.

The pressures in supersonic flow measured by the
total-pressure tube mounted on the wing surface were
found to be in close agreement with values predicted by
theory.


I NTRODUCTITON


The installation of static-pressure orifices flush
with the surface of some part of an airplane for the
measurement of pressure distribution imay not always be








NACA R3 No. LLHlOa


practicable. Static-pressure tubes and static-pressure
belts are two other means that have been used to some
extent. The validity of the pressure measurements
obtained by these means is questionable, however, because
of the possibility of effects due to misalinement with
the local air flow and, at high speeds, premature shock
formation on the static-pressure tubes and belt.

The purpose of the present investigation was to
obtain a comparison of static-pressure measurements made
by means of orifices flush with the surface, static-
pressure tubes, and a static-pressure belt in subsonic
and supersonic flow over the upper surface of an airplane
wing. As an incidental phase of the investigation, a
comparison was also obtained of the pressure measure-
ments made by means of a total-pressure tube mounted
above the wing surface outside the boundary layer with
measurements made by means of the total-pressure element
of an airspeed head mounted ahead of the airplane wing.
Measurements were made in straight flight and in turns
at airplane Mach numbers from 0.25 to 0.78 and at lift
coefficients from 0.10 to 0.68. A P-47D airplane was
used for the tests.


SYMBOLS


p static pressure

H total pressure

qc local impact pressure outside boundary layer
(Ho Pf)

x distance along chord from leading edge

a distance along surface from pressure station

c chord

!I local Mach number determined by q, and p


g acceleration of gravity








NACA R3 No. LlHlOa


Subscripts:

o free stream

f at flush orifice

b at belt orifice

t at static-pressure tube or total-pressure tube

a at airspeed head


APFARATUS


The investigation was conducted cn a section of
the left ring of a P-L7D airplane at about 65 -lercent
semispan from the plare of symmetry. The wing of the
P-T7D airplane incorporates Republic S-3 airfoil
sections, which have pressure-distribution character-
istics similar to those of the IACA CL.-series sections.
The test section was srroothed and fired oy filling and
sanding over the forward 35 percent chord on the upper
surface and over the forward 10 percent chord on the
lower surface. In the first flight, the ur:er surface
developed a crack at the leading edge of the ammunition-
corrpartment door (at 11.5 percent chord) and could not
be kept smooth and unbroken in subsequent flights.

The installation and location of the flush orifices,
the static-pressure tubes, and the static-pressure belt
on the upper surface of the wing are shown in figures 1
and 2. The flush orifices were located at 15.7, 19.1,
and 2h.8 percent chord along the center line of the test
panel. Some tests were made vith the surface contour
around the orifices at 15.7 percent chord modified
slightly by filing down the orifice and the adjacent
surface. The change in contour, which extended about
17 inches inboard and outboard of the orifices, is shown
in figure 5. The pressures measured by the flush orifices
were referenced to the static pressure measured with an
airspeed head mounted 1 chord ahead of the leading edge
near the wing tip.








NACA R. Ho. i J t-'b.


One of the static-iressure tubes tested is shown in
combination with a total-pressure tube in figure 4. This
combination was designed for use in locating the -osition
of transition from laminar to turbulent flow in the
boundary layer. In the present tests, the total-pressure
tubes of the combinations were not used. The static-
pressure tubes were 1/8 inch in outside diameter and had
six orifices equally s'aced around the periphery at
1- inches (10 tube diam) downstream iron' the remispherical
end and 35 inches upstream from the first su,-porting
bracket. The static-pressure tubes were stationed so
that their orifices were at 15.7, 19-4, and 2L.8 per-
cent chord on the upper surface of the test section.
The axes of the tubes were raised 1/l1 inch above the
wing surface except that, for some tests, the tube at
19.f percent chord was placed in contact with the surface.
The upstream or static-pressure tube 1 was 63 inches
inboard of the row of flush orifices and tubes 2 and 5
1 1
were 4- and 12- inches inboard of the flush orifices,
respectively.

The belt, which is shown in cross section in fig-
ure 2, was made up of five Saran (vinylidene chloride)
tubes 1/8 inch in outside diameter placed side by side
and cemented to the wing surface. Filler was used
between the adjoining tubes to provide a flat surface
and next to the end tubes to fair the belt into the wing
surface. A final finish was obtained by cementing fabric
over the tubes and over several inches of the wing sur-
face on either side of the belt, applying cement to the
outside of the fabric, and then sanding the belt smooth.
The belt extended from 15 percent chord on the lower
surface, around the leading -d.&e, to 55 percent chord
on the upper surface. The center of the belt was
I- inches outboard of the flush orifices. The orifices
in the static-pressure belt were placed at the same
chordwise locations as the flush orifices and the
orifices in the static-pressure tubes. The arrangement
of the belt orifices is shown in figure 2. The pres-
sures measured by the static-pressure belt i.nd the
static-pressure tubes were referenced to the pressures
measured with the flush orifices at corresponding chord-
wise locations.







NACA RB F( o. LIHlOa


Total-pressure measurements in the flow over the
wing were made with the tube shown in figure 5. The
tube was made of copper tubing 1/8 inch in outside
diameter with a wall thickness of 1/32 inch. For the
tests, the total-pressure tube was located at 19.4 per-
cent chord in place of static-pressure tube 2 and, at
this location, was set 1 inch above the wing surface
in order to clear the boundary layer for all test con-
ditions. Total-pressure measurements were obtained with
static-pressure tube 1 in place or removed. The pressure
measured by the total-pressure tube was referenced to the
total pressure measured by the airspeed head mounted ahead
of the wing near the tip.

All pressures were recorded by an NACA multiple
recording manometer.

Surface-curvature measurements were made in the
vicinity of the static-pressure tubes and the static-
pressure orifices in the wing and belt by means of a
curvature gage of the type shown in figure 6. The
distance between the legs of the curvature gage was
7- inches. The measurements are presented in figure 7
as a plot of gage deflection against distance ahead of
and behind the location of static-pressure orifices in
the wing, belt, and static-pressure tubes.


TFSTS


Tests were made in straight flight at altitudes from
12,000 to 25,000 feet at indicated airspeeds from 150
to 410 miles per hour. Tests were also made in turns
(og to -g) at an altitude of 20,000 feet at indicated
airspeeds from $10 to 375 miles per hour. The flight
TMach numbers ranged from 0.28 to 0.78 and the airplane
lift coefficients ranged from 0.10 to 0.68. The local
Mach number of the flow over the wing ranged from 0.5-
to 1.4 at the chordwise stations where the pressure
measurements were made.


PRESENTATION OF RESULTS

The results of the investigation are presented in
figures 6 to 10. In figure 8, the difference between







NACA RY INo. LLiH1La


the pressure measured by the static-pressure tubes pt
and the pressure measured by the corresponding flush
orifices pf as a fraction of the local impact pressure
at the flush orifices outside the boundary layer qf is
plotted against the local Vach number at the flush
orifices Mf. In figure 9, the difference between the
pressure measured by the belt orifices b and the
pressure measured by tho corresponding flush orifices
Pf is similarly plotted except that for x/c = 0.157 in
test 1, for which no data were obtained with the flush
orifice, the pressure of the static-pressure tube was
used as a basis for comarison. The difference between
the pressure measured by the total-pressure tube on the
wing Ht and the pressure measured by the total-rressure
element of the airspeed head Ha as a fraction of
qcf is plotted ag'i.nst .!r in figure 10. Tihe theoretical
loss in total pressure, computed by the method in refer-
ence 1, is given in figure 10 for comparison.


DISCUSSION OF RESULTS

Static-Pressure Measurements


The results shown in figure 8 indicate that, at
subsonic velocities, the pressures measured by the
static-pressure tubes were about equal to those measured
by the flush orifices at x/c = 0.157 and 2 to 5 percent
of the local impact pressure higher than the pressures
measured by the flush orifices at x/c = 0.19 and
x/c = 0.24.. In transition f'rrr subsonic to supersonic
flow, the pressures measured by the tubes relative to
those measured by the corresponding- flush orifices
appeared to decrease in all cases by 2 to 4. percent
of the local impact pressure. For static-fressure tube i,
this decrease occurred at a local Mach number slightly
greater than 1. In supersonic flow, the pressures were
generally lower for static-pressure tubes 1 and 2 but
higher for static-pressure tube 3 than the pressures
measured by the corres~n*ir51nng: flush orifices. Data at
local iMach numbers between 0.97 and 1.20 for tubes 2
and 5 were obtained only as a normal shock wave passed
over the chordv.tse stations where the measurements were
made. The position of the normal shock wave varied
across the span of the wing, however, with the result









NACA RB No. LhiIlOa


that the flush orifices and the stat!c-pressure tubes
were in different stages of a steep pressure gradient
associated with shock. The data obtained under this
condition are not included in figure C. Sone of the
values immediately below a local Vach number of 1 in
figure 8 were obtained with shock occurring upstream
of the measurement station.

The variation with chordwise location of the
differences between the pressures mrrasured by the
static-pressure tubes and the flush orifices may be
due to differences in al!inr.ent of the tubes .with
local air flow, to differences in the surface contour
at the flush orifices and the tubes (fig. 7), or to
differences in the extent to which the tubes were sub-
merged in the boundary layer. in an attempt to deter-
mine the effect of differences in surface contours, a
test was made with the surface around flush orifice 1
filed down (figs. 5 and 7). Although the results of this
test (fig. 8) were not conclusive, a tendency for the
modified orifice to measure higher pressure than the
original orifice was indicated. The thickness of the
boundary layer at 15.7, 1i.L, and 21.8 percent chord
was estimated to be about C.15, 0.25, and 0.55 inch,
respectively. These estimates were based on boundary-
layer measurements made in other tests at an inboard
station and on the assumption that transition from
laminar to turbulent flow occurred at the leading edge
of the ammunition-compartment door. In order to inves-
tigate the effect of the location of a static-pressure
tube in the boundary layer on the pressure character-
istics of the tube, a test was made with static-pressure
tube 2 placed in contact with the surface. The results,
which were obtained only in subsonic flow, show that the
static pressures measured with the tube in contact with
the surface agreed with pressures measured with the
tube 1/L inch above the surface.

A comparison in figure ? of the static pressures as
measured by the static-pressure belt and the flush
orifices shows discrepancies in some cases, particularly
for inboard belt orifice 1, between different tests made
under the same flight conditions. This effect was
probably due to the fact that the fabric which formed
the surface of the belt was not adequately cemented to
the tubes and became detached around the belt orifices
during the course of the tests. Only the results obtained
with the outboard belt orifice 1 and inboard belt
orifice 2, where this condition apparently did not occur,









NACA RB No. HlOR10a


and the results of earlier tests for the other belt
orifices should be considered as representative of the
characteristics of a suitably constructed belt. For
these cases, the difference in pressures measured by
the belt and flush orifices was less than 5 percent of
the local impact pressure and showed no large change
in the transition from subsonic to supersonic flow.
The comparison is subject to the same consideration of
the effect of surface contour at the belt and flush
orifices as in the case of the static-pressure tubes.
Data for local Mach numbers between 0.97 and 1.20 are
not included in figure 9 for reasons previously dis-
cussed.

The results in figures 3 and 9 generally indicate
that the pressure measurements obtained by means of
the static-pressure tubes and belt, if discrepancies
due to faulty belt construction are discounted, were
reasonably accurate. The critical Mach number deter-
mined by either of these methods, for example, would
probably be correct within 2 percent. These results
may not apply, however, to arrangements of static-
pressure tubes having orifices located at different
distances (in tube diam) from the nose and supporting
bracket or to static-presssure belts of greater thickness
or width in relation to the size of the wing than the
belt used in this investigation.


Total-Pressure Measurements

The pressures measured by the total-pressure tube
on the upper surface of the wing in subsonic flow were
found to agree with the pressure measured by the total-
pressure element of the airspeed head, as indicated in
figure 10. In supersonic flow over the wing, however,
the total-pressure tube on the wing measured a pressure
that was lower than the pressure measured by the air-
speed head by an amount which increased with local Mach
number. This difference in total pressures, due to the
formation of a normal shock wave just ahead of the mouth
of the tube mounted on the wing, is in close agreement
with that computed from the theory of reference 1.









NTACA RB No. L4HlOa


CONCLUSIONS


A flight investigation of several methods of
measuring static pressure and of the characteristics of
a total-pressure tube in supersonic flow has indicated
the following results:

1. The pressures measured by the static-pressure
tubes on the upper surface of the wing in subsonic flow
agreed within 5 percent of the local impact pressure
with the pressures measured by the flush orifices. In
transition from subsonic to supersonic flow, the pressures
measured by the static-pressure tube relative to those
measured by the flush orifices decreased by 2 to 4 percent
of the local impact pressure.

2. Sone results of the tests with the static-
pressure belt were influenced by effects due to faulty
construction of the belt. In other cases, however, the
pressures measured by the belt agreed within 5 percent
of the local impact pressure with the pressures measured
by the flush orifices.

5. The total-pressure tube located outside the
boundary layer on the upper surface of the wing measured
pressures in supersonic flow that were in close accord
with the values predicted by theory.


Langley memoriall Aeronautical Laboratory
National Advisory Comnittee for Aeronautics
Langley Field, Va.







REFER L 'ICE


1. Taylor, r-. I., and Maccoll, J. A..: The Mechanics of
Compressible Fluids. Tvo-Dimensional Flow at
Supersonic Speeds. Vol. III of Aerodynamic Theory,
div. H, ch. IV, sees. 2 and 5, 'F:. F. Durand, ed.,
Julius Springer (Berlin), 1935, pp. 256-242.






NACA RB No. L4H1Oa


Figure 1.- Installation of static-pressure
belt and static-pressure tubes on test
panel.


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NACA RB No. L4HlOa


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