CB No. Llt.E29 
 
 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 
 
 WARTIME REPORT 
 
 ORIGNALLY ISSUED 
 
 May 191^ as 
 Confidential Bulletin IAE29 
 
 EFFECT OF MACH NUMBER ON POSITION EKROE 
 
 AS APPLIED TO A PITOT -STATIC TUBE 
 
 LOCATED 0.55 CHOED AHEAD 
 
 OF AN AIRPLANE WING 
 
 ^y W. F. Lindfley 
 
 Langley Memorial Aeronautical Laboratory 
 Langley Field, Va. 
 
 NACA 
 
 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 
 
-71 i ^^'10^ 
 
 FACA G3 No. rJ|E29 
 
 I^ATIONAL ADVISORY CO^'ITTEE FOR AERONAUTICS 
 
 CONFIDENTIAL BULLETIN 
 
 EFFECT OP MACH NUI.IBER ON POSITION ERROR 
 A3 APPLIED TO A PITOT- STATIC TUBE 
 LOCATED 0.55 CHORD AHEAD 
 OF AN AIRPLiiNE vVING 
 By Vv. F. Llndsey 
 
 STOCvIARY 
 
 The effect on static-pressure meas-arernents of 
 locating a pitot-static tube 0.55 chord, ahead of an 
 airfoil section has been Vnvestlgatea . The tests were 
 conducted in the NACA 2li.-inch high-speed twr.iel on 
 airfoil sections of various thickness ratios over a 
 large range of l\!ach number. The results show that for 
 a v;ing having a thicl-cness ratio of O.I5 the measured 
 Each mirnber, determined from a pitot-static- tube reading, 
 Is approximately O.Gl too low at a stream ?.:ach number 
 of O.L and approximately O.OJ too low at a r.Iach number 
 of O.b. 
 
 INTRODUCTION 
 
 Pressure measurements from pitot-static tubes 
 moionted on aircraft are subject to two t^/'pes of error, 
 namely, the calibration error and the position error. 
 The calibration error is a function of the design of 
 the pitot-static t-;be, and considerable data are available 
 to shov/ the variation of this error at high Mach 
 numbers (reference 1). The position error, vvhich is 
 dependent on the location of the pitot-static tube 
 with respect to some other body, results from the 
 influence of that body on the flov/ at the pitot-static 
 tube and affects, primarily, the static pressiire. The 
 magnitude of this error and the variation of the 
 magnitude with angle of attack are usually determined 
 by low-altitude flight calibrations. The existence of 
 
2 CONFIDENTIAL NACA CB No. 1^229 
 
 a large cha.ige in position error at high Mach ntunbers 
 has b33n observed in flight Tor installations below the 
 wings of aircraft. 
 
 The data presented and discus'^.ed herein were 
 obtain3d as a part of a tunnel-wall investigation which 
 was conducted in the NaCA 24-inch high-speed tunnel and 
 in which a part of the two-dimensional-flow field 
 around an airfoil was ir.easured. The measurements of 
 the static pressure were made at tl^e tui'inel wall in the 
 region ahead of tne airfoil on the chord extended 
 (fig. IJ . Tuese measurements indicate the position 
 error In the case of many standard pitot-static-tube 
 installations but not the change in calibration error 
 resulting fron; tii:.; variations in the pitch angle of 
 the pitot-static t\ibe or the errors in the total pressure. 
 The tests were conducted for a large range of thickness 
 ratio at lift coefficients extending from approximately 
 to 0.7 and at Mach nu^nbers extending from arrroxi- 
 mately 0.25 to 0.80 
 
 APPARATUS .-.ND M2THCD 
 
 Tests of the N^^CA lC-106, NaCa 16-2:5, and NhCa 
 16-120 airfoil sections were conducted in the NaCa 
 24-inch high-speed tunnel (reference 2) with tte 
 circular test section modified by the installation 
 of flats on the tunnel walls. Thjse flats reduced 
 the span of each model from 24 to 16 inches (I'ig. 2). 
 Each model completely spanned the test section and 
 passed through holes, of the same shape but sligntly 
 larger, in large circular plates that were fitted Into 
 the flats in such a way as to rotate with the model. 
 (See fig. 2.) Pressures were measured at static- 
 pressure orifices installed in these plates ahead cf 
 the model on the chord extended, aS shown in figure 1. 
 Tnose pressure oril'ices were connected to a plioto- 
 recording multiple-tube manometer (reference 2), and 
 measurements were made for Mach numbers extending from 
 approximately 0.35 to 0.80. 
 
 RESULTS AND DISCUSSION 
 
 The symbols used in the presentation of data for 
 the present tests are as follows: 
 
 coyFiD::>TiAL 
 
N/\CA C^ 70. t1\.E2^ C 0?JP TBEITTIAL 
 
 p' static pressure ireasured at orifice ahead of 
 airfoil 
 
 p stroar;: static pressure 
 
 H total or stagnation pressure 
 
 ■I 
 
 M Mach nutnber 
 
 t/c thickneso ratio of model 
 
 a angle of attack of model 
 
 c chord of model 
 
 V-j_ indicated airspeed (true airspeed at standard sea- 
 level pressure and temperature) 
 
 h pressure altitude 
 
 The basic results presented in figure Z show that 
 the static prescijre at 0.55c ahead of the leading edge 
 of the model is greater than the stream static pres- 
 sure, in accordance with theory. This result occurs 
 because of the stagnation region that exists at the 
 leading sdge of the model. Figure 3 also E^hows tnat 
 the magnitude of the difference between the measured 
 and the stream static pressures increases as the 
 thtcknesc of the model increases because of the 
 increased extent of th3 stagnation region. 
 
 The effects of compressibility on the flow along 
 the upper and lower surfaces of an airloil have been 
 shown by experimental investigations (reference 2) and 
 by theoretical studies (ref'erences 4 to 7). No 
 quantitative ini ormation, hovvaver, has been available 
 on the flow field ahead oi" thi point of stagnation 
 pressures, that is, the region of zero velocity. 
 Because oi' the unusual condition existing at tnis 
 point, the methods which have been applied to the 
 flow along the surfaces are not believed to be appli- 
 cable to the region under consideration. 
 
 Theorcitical studies of compressible ^.'lows and 
 experimental data have shovvn, however, that the extent 
 of the region at the leading edge of models, in which 
 
 CONFIDENTIAL 
 
CONF'IDIFINTIAL NACA C3 No. 1J4.E29 
 
 the T^ressiTe Is greater thai stream pressure, increases 
 with Increasing Mach niamber. These increases are conri- 
 parable to increases In thickness, the effect of v/hlch 
 at a constant Mach uumhsr can be seen in figure 3. 
 This effect may be modified by the decrease in the 
 extent of the field of influence uDStream of the model 
 with increasln£^ Mach momber, vi^hich results because the 
 velocity of pressure propagation is equal to the 
 velocity of sound. Although no quantitative comparison 
 of the relative r.agnitudes of the tv/o effects can be 
 given from previous investigations, it can be seen in 
 figure 3 that for the present investigation the over-all 
 effect of compressibility on the field of flow is to 
 increase the pressure and thereby increase the position 
 error with increased Mach number. 
 
 The location of the pressure orifice relative to 
 the model remained unchanged for changes in angle of 
 attack. The eflect of changes in angle of attack or 
 In lilt coefficient on the static-pressure error is 
 small within the lihiits of this investigation. A 
 more extensive investigation of the field of flow 
 ahead of the model would be needed to explain the 
 reason for thj difference in direction of the cl-ange 
 in this static-pressure error with increase in lift 
 coefficient at M = 0.4 for the airfoils having 
 thickness ratios of 0.06 and 0.30. (See fig. 3.) 
 
 The effect of pressure altitude on the variation 
 of the Dosition error with indicated airspeed is 
 presented in figure 4. These variations with altitude 
 are a result of the variations of r.'ach number with 
 altitude at a constant Indicated airspeed. 
 
 The basic results of figure 3 for an angle of 
 attack of C*-* have been converted to show, in figure 5, 
 the error in Mach momber resulting from the increased 
 static pressure. Ti.e m.ost significant result indicated 
 in this figure is tiie effect of position error on the 
 magnitude of the erior in the determination of the I.'ach 
 number. For the airfoil v;ith a thickness ratio of 0.15, 
 the J.iach ntur.ber decrement is approximately 0.01 at a 
 Iwach number of 0.4 and 0.03 at a Mach nui:iber of 0.8; 
 for the airfoil with t/c = O.oO, the Mach number 
 
 COI'.'FIDIUTTAL 
 
NACA CP No. 14^2 9 CONFIE'SNTIAL 
 
 decrement is approxi^^^atel^r 0.01 at a Mach number 
 
 of 0.25 and O.bS at a Mach number of O.72. The 
 
 decrement for airfoils having a thlcjcness ratio 
 
 of 0.06 or ]ess is nrobablj'" within the usual accuracy 
 
 of measurements. 
 
 Langlev Memorial Aeronautical Laboratory, 
 
 National Advisorv Co-rimittee for Aeronautics, 
 Langley Field, Va . 
 
 REFERENCES 
 
 1. Hensley, Reece V.: Calibrations of Pitct-Static 
 
 Tubes at High Speeds. N.-.GA ACR, July 19li-2. 
 
 2. Stf.ck, John, Linrlsej, IV. F., and Littell, Robert E. 
 
 The Com-oresslbility Burble and the Effect of 
 Com-DressibilltT en Pressures and Forces iiCtlng 
 on an Airfoil. KAGA Rep. No. 6)4.6, I938. 
 
 3. Pinkerton, Robert T,'.: Calc^^lated and I.'easured 
 
 pressure Distributions over the Uidsnan Section 
 of the N.A.C.A. 1|1).12 Airfoil. NACA Reo. 
 No. 565, 1956. 
 
 Ij.. Glauert, H.: The Effect of Comr^resslbility on the 
 Lift of an Aerofoil. R. 8c l{. No. 1135, 
 British A.R.G., 1927. 
 
 ■3 
 
 Prandtl, L. : General Considerations on the Flow of 
 Comm-'esslble Fluids. NfiCA TM No. S05, I936. 
 
 6. von Karman, Th . : Compressibility Effects in Aero- 
 
 d^mam.ics. Jour. Aero. Sci., vol. 3, no. 9» 
 July 191^1, pp. 337-55^^. 
 
 7. Kaplan, Carl: The Flow of a Compressible Fluid 
 
 past a Curved Surface. NACA aRR No. 3K02, 19i4.3. 
 
 CONFIDENTIAL 
 
Digitized by tlie Internet Arcliive 
 
 in 2011 witli funding from 
 
 University of Florida, George A. Smathers Libraries with support from LYRASIS and the Sloan Foundation 
 
 http://www.archive.org/details/effectofmachnumbOOIang 
 
NACA CB No. L4E29 
 
 CONFIDENTIAL 
 
 Fig. 1 
 
 Static pressure 
 orifice iocafed 
 0.55c o lie ad of 
 
 mod^l ieading 
 
 e>dge 
 
 Orig/na/ 
 \ v^alls 
 
 ^ Pre'senf Wails 
 y\ farmed by f/a/s 
 
 /^ 
 
 S-inch ci^ord 
 /vodei 
 
 F/ane of 
 CO a bra fed 
 sfa tic pressure 
 orif/ces 
 
 NATIONAL ADVISORY 
 COMMITTEE FOR AERONAUTICS 
 
 figure /. - Modi fled test section of A/AC A 
 ^4-mch /7/g/7 -speed iunne/. 
 
 CONFIDENTIAL 
 
NACA CB No. L4E29 
 
 CONFIDENTIAL 
 
 Fig. 2 
 
 (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-inch 
 
 high-speed tunnel- 
 
 CONFIDENTIAL 
 
NACA CB No. L4E29 
 
 Fig. 3 
 
NACA CB No. L4E29 
 
 Fip, 
 
NACA CB No. L4E29 
 
 Fig. 5 
 
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