Effect of Mach number on position error as applied to a pitot-static tube located 0.55 chord ahead of an airplane wing

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Material Information

Title:
Effect of Mach number on position error as applied to a pitot-static tube located 0.55 chord ahead of an airplane wing
Alternate Title:
NACA wartime reports
Physical Description:
5, 5 p. : ill. ; 28 cm.
Language:
English
Creator:
Lindsey, W. F
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:
Airplanes -- Wings   ( lcsh )
Aerofoils   ( lcsh )
Aerodynamics -- Research   ( lcsh )
Genre:
federal government publication   ( marcgt )
bibliography   ( marcgt )
technical report   ( marcgt )
non-fiction   ( marcgt )

Notes

Summary:
Summary: The effect on static-pressure measurements of locating a pitot-static tube 0.55 chord ahead of an airfoil section has been investigated. The tests were conducted in the NACA 24-inch high-speed tunnel on airfoil sections of various thickness ratios over a large range of Mach number. The results show that for a wing having a thickness ratio of 0.15 the measured Mach number, determined from a pitot-static-tube reading, is approximately 0.01 too low at a stream Mach number of 0.4 and approximately 0.03 too low at a Mach number of 0.8.
Bibliography:
Includes bibliographic references (p. 5).
Statement of Responsibility:
by W.F. Lindsey.
General Note:
"Report no. L-75."
General Note:
"Originally issued May 1944 as Confidential Bulletin L4E29."
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 - 003638560
oclc - 71835705
sobekcm - AA00006293_00001
System ID:
AA00006293:00001

Full Text
1


CB No. L4E29


NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS


WARIRTIMEI REPORT
ORIGINALLY ISSUED
May 1944 as
Confidential Bulletin L4E29

EFFECT OF MACH NUMBER ON POSITION ERROR
AS APPLIED TO A PITOT-STATIC TUBE
LOCATED 0.55 CHORD AHEAD
OF AN AIRPLANE WIG
By W. F. Lindsey

Langley Memorial Aeronautical Laboratory
Langley Field., Va.





S: .. : .... .



WASHINGTON
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 pre-
viously 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 75


DOCUMENTS DEPARTMENT


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' -.3 0'o. LirE29


NA'ICAL AVi'.'TSORY CO'TM:ITTE'E 7' A.ERONAUTICS


T,2;" TU-iTT 'T L 3.1T,--LT 7 I


-FFECT OF :.!ACH I.UT.BER 0:' FLITION ERROR

AS APPLIED TO A PITOT-3TATIC TUSE

LOCATED 0.55 CHORD A'EAD

OF ..N 't IFl T ,l_-, :E ."'ITlG

By 'V. F. Lindsey


STJTrL'ARY


The effect on t ti -rr s'vre '.e.'IsjreI ents of
locating a pitot-stLaic t"be u. 5 ch! .r ie1 C an
air.oll section ha- been '"r:v-st:irste T'e tlsts .:'ere
conducted in the .IACA 22- nch h h-speelj :.-.-1el rn
airfoil c3ctiv ." ; -ar'i':s thick ies ra f: T over a
lar.pe range -*f 1'ac.n rinumb r The result' hc'.- that for
a wvir. having a th.ic'.:ness r-'tio of 0.15 tl-e measured
Mach number, Aeter-ined from a pitot-ztatic-t'.be reading,
is anrroximatel:' J.'1. tco lo' st a stream .''a;h number
of 0. and approximately 0.03 too low at a "ach number
of 0._.


I DT.- CDU L' IC i


Pressure measur:rents fromrn pitot-static tub:-s
mocinted on aircraft are subject to two tyPes of error,
namely, the calibration error ,in9 the position error.
The calibration error is a function of the design cf
the pitot-st.-tic t-ibe, Jand considerable data are available
to show the variation of this error at high Mach
numbers (reference 1). THe position error, which h is
dependent on the location of the pitor-static tube
with respect to so-ne cth'er bodv, results froq the
influence of that body on the flow at the pitt-3static
tube and affects, primarily, Lhe static rres3uire. The
magnitude of this error and the variation of the
magnitude with angle of sttaz': are usually determined
by low-altitude flight calibrations. The existence of








2 C O::0.,T -Y" T L U:.'A CB ::n. T';29


a lar.; change in position error at high Mach numbers
has b33n observed in fl.iht for installations below the
wl-.-s of aircraft.

The data presented and discus-vc herein were
obtainAd as a part of a tunnel-wall investi-ation i':hich
was conducted in the NACA 24-inch h.;--. -d tunnel and
in which a -.'iit of the two-diannsional-flow field
arour. an airifil was measure.' i'., measurements of
the static pressure were made at the tunnel -,all in the
re -on ahead of the airfoil on the chord exterd.d
(fig. 1). Tnese measurements indicate the -c.zition
error in the casa of .:i.o7 standard pitot-static-tube
installations but not the c:'-...r 9- in calibration error
resulti, from th.li variations in the pitch a:. -l of
the pitot-static tube or the crrors in the total pressure.
I" .. tests were conducted for a la? rans of thickness
ratio at 1'ft coefficients exte. 3'-' .- r' a 'x ...-tely 0
to 0.7 and at Mach nu.nbers extending f;'oi., a-proxi-
mately 0.35 to 0.90


2AP r "'S .-D ".OD


Tests of tbo "c."iA 10-1C, NACA 16-...5, and ':rCM
16-120 airfoil sections were con.it-ted in the ::.'..C
24-inch 1.Kili-speed tunnel (reference 2) with tth
circular test section modified '- tLe installation
of flats on the tunnel walls. Trjse flats recuc7d
the span of each rmodl f.?.'. C4 to 18 inches (fig. 2).
-Tu model compile, ly s'- ..--d the test section rn,
.*sed thro'u.: h holes, o' the cameT s.:..-j but sl :.tly
lAi:-er, in large circular plates that were fitted Into
the flats in such a way as to rotate with the model.
(.a it,-. 2.) F:,ssures ware n.aasured at static-
or.esuru orl.:ices installed in thIse plates 1--iJA of
the model. on th chord xt. .-, as shown 1-, fi -ur; 1.
T:..se pressure '"rfices were con3ected to a ; ito-
recor '1:.- multi le-tube: .r ,i'.lter (ref renco 3), .:r,
measurements were mais for Mpach numbers e..t.i-'l. :: fi n
a' !roximately 0.35 to 0.,.


,:W3ULPS T.'-^ D:3,U5IC"


:-.- symbols ;.". in the resen'ation of d"i.ti for
the present tests are as follows.








rA1CA C, 7'o. E729? CO('FTDEI'TIAL


p' static pressure measuredd at orifice ahead of
airfoil

r streai'n static t;res-ure

H t.-ttl or .ta:n.ation pressure

Cz section lift co6fficient

J iM.ach nua jr

t/c thicknes: ratio of mcdel

a an.le of attack of nodel

c chcrd of nimcel

Vi indicated airspeed (true airsoped at standard sea-
level pressure and tam erature)

h pr';scure altitude

The 1:.asic results pres.ente,. in figureo show that
the' static r-3so:r-e .t ,O.5bc aiea;d of tn? l-eadini e drs
of the model is gre at' er tlan ti- Ctr-eia: static rees-
sure, in accordar.2e with th'.:."y. Th:.s 'zsult occurs
tb cause nf the sta,'natio.- r-e.rjr. t-:at exItts at the
leadin? dj e :f tin- .., l. Fl ure .so chow: t.at
the msan tude of th: dif ,ri tetv', =et n 7t. aasCur.--d
and the strea:1 static cre4sures increases the
thickness of t'.e model increasses e cause or the
increased extent of th sta.nati?.: i-e.ion.

The 3:f'scts of corn.pr.easibillity on the flow alDn g
the urper .ad lcwer surfaces cf a.i aricoil ha"e b-aen
shown by exrerimental invcsti Lg tions (reT'erance '2) :.:.d
tby th -retical studius referencess 4 to 7). o
quant Itat ive inrir::. tin, ho evr: i Las beL! avi ilabl 1
on t; h :1.w fildi ahead :,i th. ooi:nt of s-am-natic.1
pressurZ2 s, that is, tlj, ; '...-. of, z-ro :-l"ci .
r3csuuse oi' the unusual cor.nclticrn iscti: -l e t th:is
P:i:-t, ti:e methods -whiic' havO teen in.' li:a to tr.e
f1ow along thj surfaces are art i.lielad to be .-ppli-
cable to the r.'31in under c-:n- deir .ron.
Th-orejtical studies of c-esie and
experimental data ha':e shwcvn, ho-evr, :'t t. ext;nt
cf the r ron at the leading ed.= of' r.:od i.s n wrli_.


?C":FI5T .'T ,1T.L








4 CC"-'T- -TTAL !ACA CB No. L4E29


the cressrre is ---,ater tlani stream pressure, increases
..th increasing Mach number. I.-se increases are com-
parable to increases in thickness, the effect of which
at a constant ::ech urnmo.r can be seen in figure 3.
T.is effect may be modified by th. decrease in the
extent of the field of influence upstream of the model
with increas':L' :.a.h number, which results because the
velocity of pressure proragation is equ'l to the
velocity of sound. Although no quantitative c:_-. arson
of the relative r a-.tudes of the two effects can be
-iven f;-.:. previous investi.- .tions, it can be seen in
fil.'.' 3 that fcr the present investi -.tion the over-all
ef'dect of compressibility on ths field of flow is to
increase. the pressure and thIrsr. increase the ".-.:'ition
error with increased L..ich number.

"i_! location of the pressure orifice relative to
the m:,ajil remained uL--.-hned for chn:-jr-s in a:-le of
attack. TL effect of ch} n-:-s in -..:le of attack or
in lift coefficient on the static-pressure error is
small within the limits of this invest'-,i::tion. A
more extensive investi--aiton of the field of flow
ahead of the model would be needed to explain the
reason for the ;ffere nce in direction of the cL-an-e
in this static-pressure error -ith increase in lift
coefficient at .' = 0.4 for the airfoils havi:.g
thickness ratios of 0.06 and 0.30. (?e:- f-. 3.)

:-.3 effect of pressure altitude on t'-. variation
of the oosition error with I ,iicated airrs-'d is
presented in figure 4. T',.se variations with altitie
are a result of the variations of ':ach number ,'.'th
altitude at a constant indicated airs:--.;

'ir- basic results of 'irure 3 for an <--rl of
attack of Co have been converted to show, in ft._.-'- 5,
the error in "..'ch iin:.t3r result .i;g ircr., the incre: '.ed
static pressure. .: most sl ..iicant result ...-ccated
in t' 1- fl- .. is the effect of -osition error on the
nr::-":tuJ- of tl.e error in t'e determination of the ;:~ch
number. For the airfoil with a thickness ratio of 0.15,
the "--ch number decrement is a -i :::r..atel" 0.01 at a
-.ach number of 0.4 .:-. J.J at a ...a:. nu.:ber 2: .t
for the airfoil with t/c = 0.cO, the ::.ch num-rer








I,'A;A CP No. LJ429


decre-ent is aonroxi-atel- 0.01 st a Mach number
of 0.25 and 0.06 at a Yach number of 0.72. The
decrement for airfoils having a thickness ratio
of 0.06 or less -s rjrobabl"1 within the usual accuracy
of measurements.


Langleyv 7'emorial Aeronautical Laborator",
rational Advisor'. Cc'.mittae for -.eronautics,
Langler Field, .a.






1. Iensle c .: Callibraticns of Pitct-3t-t ti
Tub'g)s at gh Srp.e:' 3. :.CAr .-C7-., July 14" .

2. Stock, Jchn, LilIser, ... P., n.1 Littell, Robert E.:
.- Cc, z, -'s. iO' i Tor 'e A.rd tze 'feet of
0o- t 1 s Z '. l i .; C r .-.-* : '._. S E tin
on at- Air:oil. iFJ.. :.Ra Io. i5,. 1 r)..

5. Pir,!.-rton, 1 tert ".: ?alcu'i:.ted and 'esaLr-.d
F re s,..re .'T stL-i utior.ns .v r t.hL .:i:.-. an Section
of t-e N.! .C! ... L412 ..irfoil. -"i' -'en.
Tio. 565, 125'..

S. Glauert, H.-: T:-: Effect of ,c.rnroes ibility on the.
Lift of an Aerofoil. 2. ;. 1o. 1_55,
Sritish A.; .. ., le:7.

5 -. Prndti, L.: O-niral ros i'jbtrationr.s on the Flow of
Con-ressible "iis 0. T1 c. S905 106.

6. von !K'Er'-, n, Tr.: Corrr--ss bilit.' Effects in i'ero-
d'yn7_mics. Jour. Ae:'-. 3c.., vol. nc. ,
Jul: 1!9 1, pp. 357- .

7. Kaplan, Carl: Th-b Flo!,' of a Co~~-'essible Fluid
rast Curved SJ rfiace ?I-CA .1.2; "io. $'':02, 19.3.



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NACA CB No. L4E29


Ori 9/nll
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Present wo//s
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Sfatic pressure
orifice located
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model /ead/g9
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P/ane of
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orifc es


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NATIONAL ADVISORY
COMMITTEE FOR AERONAUTICS


Figure /. A4odifie& test section of /VAC/A
E4-mnch h gh-speed tunne/.


CONFIDENTIAL


CONFIDENTIAL


Fig. 1










NACA CB No. L4E29


(a) Over-all view with access door removed showing
model installation.


(b) Downstream view with model in place.


Figure 2.- Modified test section of the NACA 24-irch
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CONFIDENTIAL


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