RB No. I^HlOa NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WARTIME REPORT ORIGINALLY ISSUED November 19^U as Restricted Bulletin iAELOa FLIGHT INVESTIGATION AT HIGH MACH NUMBERS OF SEVERAL METH0I6 OF MEASURING STATIC PRESSURE ON AN AIRPLANE WING By John A. Zalovcik and Fred L. Daum 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 - 90 DOCUMENTS DEPARTMENT Digitized by the Internet Archive 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 7»i f*1 %* NACA RB No. LltflOa NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS RESTRICTED BULLETIN FLIOHT INVESTIGATION AT HIGH MACH NUMBERS OF SEVERAL METHODS OF MEASURING STATIC PRESSURE ON AN AIRPLANE WING By John A. Zalovcik and Fred L. Daum SUMMARt 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 P-L.7D airplane over a range of flight conditions in which local Mach numbers from O.3I1 to l.Ij.1 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. INTRODUCTION The installation of static-pressure orifices flush with the surface of some part of an airplane for the measurement of pressure distribution may not always be NACA RB No. iiiHlOa 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 O.78 and at lift coefficients from 0.10 to 0.68. A 1-1+7D airplane was used for the tests. SYMBOLS p static pressure H total pressure q local impact pressure outside boundary layer ( H o ' Pf) x distance along chord from leading edge s distance along surface from pressure station c chord M local Mach number determined by q c and p g acceleration of gravity NASA R3 No. iijHlOa 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 APPARATUS The investigation was conducted en a section of the left wing of a T-L r J'D airplane at about 63 percent semi span from the plane of symmetry. The wing of the P-1±7D airplane incorporates Republic S-3 airfoil sections, which have pressure-distribution character- istics similar to those of the KACA 23C-aeries sections. The test section was smoothed and faired by 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 upper surface develooed a crack at the leading edge of the ammuniticn- compartment 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-'-> and 2U.0 percent chord along the center line of the test panel. Some tests were made with 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 If- inches inboard and outboard of the orifices, is shown in figure 3- 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. h HACA : >. 3 .. 10a One of the static-pressure tubes tested is shown in combination with a total-pressure tube in figure i|. This combination was designed for use in locating the position 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 1/3 Inch in outside diameter and had six orifices equally spaaed around the periphery at lp inches (10 tube diam) downstream from the hemispherical o end and Jf- inches upstream from the first supporting U bracket. The static-pressure tubes were stationed so that their orifices were at 15-7» 19«U> &&& 2L.8 per- cent chord on the upper surface of the test section. The axes of the tubes were raised l/l inch above the wing surface except that, for some tests, the tube at 19. a percent chord was placed in contact with the surface. - 3 The upstream or static-pressure tube 1 was by- inches inboard of the row of flush orifices and tubes 2 and 3 were h r y and 1-- 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 bolt extended from 15 percent chord on the lower surface, around the leading edge, to 35 percent chord on the upper surface. The center of the belt was L— 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 and the static-pressure tubes were referenced to the pressures measured with the flush orifices at corresponding chord- wise locations. liiHlOa 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 lee's of the curvature gage was 1 , 7p- 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. TESTS Tests were irade in straight flight at altitudes from 12,000 to 25,000 feet at indicated airspeeds from 150 to iilO miles per hour. Tests were also made in turns fljg to 4^-gJ at an altitude of 20,000 feet at indicated airspeeds from 3 10 to 375 miles per hour. The flight Maeh numbers ranged from C.2S to O.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 O.34 to l.ij.1 at the chordwise stations where the pressure measurements were made. PRESENTATION OF RESULTS The results of the investigation are presented in figures 8 to 10. In figure 8, the difference between NACA RB No. iiiHlOa the pressure measured by the static-pressure tubes p. and the pressure measured by the corresponding flush- orifices p f as a fraction of the local impact pressure at the flush orifices outside the boundary layer q is T plotted against the local Mach number at the flush orifices Mf. In figure 9> the difference between the pressure measured by the belt orifices p^ and the pressure measured by the corresponding flush orifices p f 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 comparison. The difference between the pressure measured by the total-pressure tube on the wing Rj- and the pressure measured by the total-pressure element of the airspeed head H a as a fraction of q c is plotted against Mf in figure 10. The theoretical loss in total pressure, computed by the method in refer- ence 1, is given in figure 10 for comparison. DISCISSION OP RESULTS Static -Pressure Measurements The results shown in figure 3 indicate that, at subsonic velocities, the pressures measured by the static-pressure tubes were about equal to these measured by the flush orifices at x/c = 0.157 an ^ 2 to 3 percent of the local impact pressure higher than the pressures measured by the flush orifices at x/c = O.I9L. and x/c = 0.2)46. In transition from 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 h percent of the local impact pressure. For static-pressure tube 1, 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 corresponding flush orifices. Data at local Mach numbers between 0. Q 7 and 1.20 for tubes 2 and 3 were obtained only as a normal shock wave passed over the chordwlse stations where the measurements were made. The position of the normal shock wave varied across the span of the wing, however, with the result. NASA RB No. iiiJUOa that the flush orifices and the statj c -press ure tubes were in different stages of a steep pressure gradient associated with shock. The data obtained under this condition are not included in figure 8. Some of the values immediately below a local Mach 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 measured by the static-pressure tubes and the flush orifices may be due to differences in alinement cf 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. $ 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. I" 1 -L, and 2I4..8 percent chord was estimated to be about 0.15, 0-25, an ^- ^«35 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 doer. 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 wi th the tube l/[i. inch above the surface. A comparison in figure 9 °? 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, 8 NACA RB No. LLjHIOa 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 3 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 C.97 and 1.20 are not included in figure 9 for reasons previously dis- cussed. The results in figures 8 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 ?"ach 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-pressure belts of greater thickness or width in relation to the size of the wing than the belt used in this investigation. Total-Fressure 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. NACA R3 No. LijHlOa CONCLUSIONS A flight investigation of several methods of measuring static pressure and of the characteristics of a tobal-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 3 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 l\. percent of the local impact pressure. 2. Some 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. 3. 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 Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics langley Field, Va . REFERENCE 1. Taylor, C-. I., and Neccoll, J. ".'..: The Kechanics of Compressible Fluids. Two-Dimensional Flow at Supersonic Speeds. Vol. Ill of Aerodynamic Theory, div. E, ch. IV, sees. 2 and 3> '■'■'. F. Durand, ed., Julius Springer (Berlin), 1935, PP. 236-2ij.2. NACA RB No. L4Hl0a Fig. 1 Figure 1.- Installation of static-pressure belt and static-pressure tubes on test panel. NACA RB No. L4Hl0a Fig, Q) ■-4_ i- '-+-. £ -2 o£ ^3 "2 ^ O -Q -ii C +- L^ O c o 4- CJ i_ ■+- 4- JC CD (/; Q J5 <+_ CD U- 15 o (/J cO c Q) o s_ 4- O- o ( ) i ; o 4— __J D 1 CO t\i CD V- 3 CT U- 'ig. NACA RB No. L4H10a ? \J • o II I ) < 1 , J V T iO 10 en 11 o O uj o iO o £ LQ -C 01 C > O a Q) d) 3 CO CO
  • u 3 O I (£1 c/c=0 .194. 1.5 1.0 .5 .5 1.0 1.5 5, in. - .16 .14 a- o u> .12 ; >1 r ^ X b? fc^ ^ fccg=< r^ !>o^ ^5^ -<— end inq e dqe_ ft' t/c=0. 24*. 1.5 1.0 .5 j, in. 1.0 1.5 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Rqun: 7- Surface-curvature measurements near pressure stations. Fig. 8 NACA RB No. L4H10a £0 r. W i,° V O -o-tff ■d* JV & £ (To cr (o) Static -pressure tube J- x^: =0-157. Test o 1 o 4 o 5 a 6 Static - pressure tube £in contact with surface ? 7 Static- pressure tube £ in contact with surface v 8 Flush orifice 1 modified ^ 9 6^^y Downstream of shocK — ' ■ ' — . — i i. — i — , — \ 1 — i i 1 — (b) Static -pressure tube 2 ; x/c -0.I94 .1 £° 7 %& 7 .8 1.0 I.Z Mi (c) Static- pressure tube 3-,y.Jb=0ZA3- NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS 1.4- Figure 8.- Comparison of pressures measured by stai pressure tubes and flush orifices. NACA RB No. L4Hl0a Fig. 9 _ Test _ ITJL -5 LJ— 3 &~n + o o f^— t-2 '""' -% 1 i o n ii u X g (a) Inboard orifice 1 ; -y/c=QI5"7. i o ^ Ji < 1-n (b; Out boa rd onf ice S -v— **■ t- J, tl/c* 0.157 ^ Test + 2 .X 3 a 4 U.l <3- I Test O 5 Data for inboard orifice 1 ^ 6 pbtted os[p-p\lq r ifsflu — v 7 t7 8-> F| ush orifice i •<3 9 j modified s^ Down-stream of 5hocK. ^j= r -j. r T - ^ (Cj Inboard orifice £-., ^/c=Ql94 J * o ^^r^r ^ r T 1 o ? V 7 ^ (d; Outboard orifice 2 ; y/c=O.I94. o .1 .a? n 3 O n O -^ s-\- **. o -j^Sd o 4 .8 1.0 M, AZ A4- (e) Orifice 3 } V C= O-^ 4 ®- NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS Fiqare 9.- Comparison of pressures measured by static -pressure belt and flush orifices. Fig. 10 NACA RB No. L4H10a -o 0) o b Q_ a> o a) -Q c -G ^ 3 -o b -*-> « ? V ^ => 1_ *> ^ «»— (O "> o 0) 4J L. ^ TJ Q_ Q- O 1 • J^ o o £ CO V) ~o o €N ^ -4-> 00 0) 01 L. o -•-> -*j w i0 tfJ -C