Drag Measurements at Transonic Speeds of NACA 65-009 Airfoils Mounted on a Freely Falling Body to Determine the Effects of Sweepback and Aspect Ratio

Copy No. ::4

CONFIDENTIAL

RM No. L6KO8c

FILES OF 4IS DOCUMENT ON LOAN FROM THE 'MMT11 FOR AERONAUTICS Py

NACA

Lit A)RESSED

RESEARCH AUT-MEMORANDUM

CACEr,LED T:T BY AUT:: y J. 7.clic

DRAG MEASUREMENTS AT TRANSONIC SPEEDS OF

NACA 65-009 AIRFOILS MOUNTED ON A FREELY

FALLING BODY TO DETERMINE THE EFFECTS

OF SWEEPBACK AND ASPECT RATIO

By

Charles W. Mathews and Jim Rogers Thompson

Langley Memorial Aeronautical Laboratory

Langley Field, Va.

Restriction/Classification Cancelled revilatlot of its — 1 . In any manner N an n,lno,1nadporte,, In prrhtblte by inn.

tnf,rmytloo so ilnefinil .s,,y be Impantod olvly I: ernons in the military ned naval ten-v-ly e of It,, a!ted theta,,, 4,P—PMete vivlflvov IfiIorO and employees of the Federal (IOvortoe yet -elm bane a lvvgltt,eave let-rent therein, ted to 'JaNsd State. niliaene of tn-two loyalty and thayrstirn eN' at necesntty maya be Inforenad thea-n1.

NATIONAL ADVISORY COMMITTEE

FOR AERONAUTICS

WASHINGTON

January 22, 1947

(o'

Restriction/Classification Cancelled MACA RM No, L6KO8c NATIONAL AD1JISOY COMMITTEE FOR AERONAUTICS ESEARc: MEL JORANDUM DRAG MEASUREMENTS AT TRANSONIC SPEEDS OF NACA 65-009 AIFOIL$ MOTJTED ON A FREELY FALLING BODY TO DETERMINE THE EFFECTS OF SWEEPBACK AND ASPECT RATIO

By Charles W. Mathews and Jim Rogers Thompson

SUMMARY Drag measurements at transonic speeds on rectangular airfoils and on airfoils swept back are reported.

These airfoils, which were munted on cylindrical test bodies,are part of a series being tested in free drops from high altitude to determine the effect of variation of basic airfoil parameters on airfoil drag characteristics at transonic speeds. These rectangular and swept-back airfoils had the same span, airfoil section (N-ACA 65-009), and chord per pendicular to the leeding edge. The tests were made to com p are the drag of rectangular and swept- back airfoils at a higher aspect ratio than had been used in a similar comparison reported previously.

The results showed that the drag of the swept-back airfoil was less than 0.15 that of the rectangular airfoil at a Mach number of 1.00 and less than 0.30 that of the rectangular airfoil at a Mach number of 1.17.

A comparison of these swept-back airfoils with similar airfoils of lower aspect ratio previously tested by the same method indicated that in the investigated speed range reduction in aspect ratio results in increased drag. In the highest part of the investigated speed range, however, the drag coefficient of the high-aspect-ratio swept-back airfoils showed a tendency to approach that of the lower-aspect-ratio swe p t-back airfoils. A similar comp. son for the rectangular airfoils showed - Restriction/Classification Cancelled

CFIDE;rTIAL NACA Rt -Co. L6I(O8 that dela y in the drag rise and a reduction in drag at supercritical speeds can be realIzed through reduction ratio.

in aspect These results confirm those reported In i\bC iCH No. L5J16.

,INPFtUDUCT10N A serious limitation on practical flight In the transonic-speed. range results from the large abrupt increases in drag of conventional- airplane configurations as sonic speed is approached. Because of the impor tance of this problem, a series of tests is being conducted at the Lan1ey Memorial Aeronautical Laboratory of the NACA to determine aerodynamic shapes and configurations that have a --inimlan of dra g at transonic speeds. In these tests, data are teler' etered from special test configurations durinr, free fs3.1 from high altitude. Previous

tests

In which this method was employed ieie reported in references 1 and. 2. The object of the present tests was to compare the drag of rectangular and swept-hack airfoils at a higher aspect ratio than had been used a similar conoarisori reported n reference 2.

For the tests reported herein dra g measurements ware made on rectangular airfoils and on airfoils having 45° sweepback.

These airfoils incor p erated NAC 65-009 sections of equal chord perneridicular to the leading edge and d.ffersd from the airfoils of reference 2 only 'by an increase in s p an.

The subscri p t 1 has been deleted from the desicnatjon of NktCA Sserles airfoils with thickness ratios less t.aani 0.12 rf thechord. The airfoil designated 65-009 in the present paper, therefore, Is the airfoil

section

designated 65-009 in reference 2.

Th results of the tests on these airfoils are pre-

sented as

curves showing the variation of drag cceffi- dent with Mach number in the transonic-s p e range.

Comp arab curves are also presented for the airfoils of reference 2.

CONFIDNTIAL

NACA RM WOe L6K080 CONFIDENTIAL APPARATUS AND METHOD Test configurations.- The general arrangement of the two test configurations is shown in figure 1, and details and dimensions are shown in figure 2 The airfoils were mounted on bodies identical wi.th those of reference 2.

3 -inch diameter and were These bodies were cylinders of 10 fitted with a pointed nose and a small tail fairing.

The airfoils were. located near the rear of the cylin- drical part of the body and entered the body through rectangular slots gLinches long and 1 inch wide. The airfoils were staggered so that front and.rear airfoil sets could be mounted on separate spring balances. This arrangement has the additional advantage of reducing the interference effects of the rear airfoil on the front airfoil.

Both the rectangular aIrfoils and the airfoils having 65-009 sections of 0-inch chord sweepback had NACA 45 0 perpendicular to the leading edge and equal spans of 25 inches outboard of the body as compared with 15 inches for the otherwise identical airfoils of reference 2.

The nominal aspect ratio b 2/S for the present 3.6 for the swept-back airfoils was 5.4 as compared with swept-back airfoils previously tested, where b is the over-all span of each airfoil set and S is the plan area of each airfoil set including that within the body.

Corresponding nominal aspect ratios for the rectangular r7.6 and 5.1, respectively.

airfoils were Measurements.- Measurement of the desired quantities was accomplished as in the previous tests (reference 2) through use of the NACA radio-telemetering system and radar and phototheodolite equipment. The following quantities were recorded at two separate ground stations by the telemetering system: (1) The force exerted on the body by each set of airfoils as measured by a spring balance (2) The total retardation of the body and airfoils as measured by a sensitive accelerometer aimed with the longitudinal axis of the body CONFIDENTIAL

CONFIDENTIAL NACA RM No. L6KO8c The local static pressure at a station on the (3) body 11 chords ahead of the front airfoil as measured by four orifices connected to i'aneroid pressure cell (see fig. 2).

A time history of the position of the body during its fall was recorded with respect to ground axes by the radar and phototheodolite equipment, and a survey of atmospheric conditions applying to each test was obtained from synchronized records of atmospheric pressure, tem- perature, and geometric altitude taken during the descent of the airplane from which the bodies were dropped.

Reduction of data.- As in the previous tests the velocity V of the body during its fall was obtained both by differentiation of the flight path as determined from radar and phototheodolite records and by integra- tion of the vector sums of gravitational acceleration and the directed retardation as measured by the acceler- ometer. The drag D of each set of airfoils was obtained from the relation D = R + Wtae where R measured reaction between airfoil and body, pounds W weight of airfoil assembly, pounds reading. of accelerometer (retardation), g ae The atmospheric pressure p, . the temperature T; and the airfoil frontal area F were combined with the simultaneous values of velocity and airfoil drag to obtain Mach number M and the ratio D/Fp. A curve of this parameter D/Fp against Mach number affords a simple and convenient means for expressing drag in the transonic-speed range as a function of Mach number, .

altitude, and size, Values of conventional drag coefficient based on frontal area CD were obtained from the. relation F . .

• _D/Fp Y CONFIDENTIAL

pro.

n AA L6:o8 Coi'rIDE:TIAL where the ratio of specific heats was taken as 1.4.

y Drag coefficient based on plan area was obtained by CD multiplying values of by theratio of frontal area CD to plan area. Areas used did not include those within the body.

RESULTS AO - ,'j DISCUSSION: • Time histories of the pertinent quantities obtained from each test are given in figures 3 and 4. A check on the over-all accuracy of the velocity and retardation measurements is provided in these figures by a comparison of the velocity obtained from flight-path data (test points and solid fairing) with the velocity obtained from acceleratxn data (dashed fairing). The maximum discrep- ancy in velocity obtained by these two independent methods may be seen to be about 10 miles per hour. If the srurce of this error is wholly in the measurement of retardation, the corresponding mean accelerometer error would be of the order of 0.01g, which is within the expected limit of accuracy -, f this instrument. The velocities used to compute the Mach number were those taken from the fairing of the flight-path data. The acceleration data, however, were used as a guide-for this fairing, particularly in the fairing of the last few seconds of the data shown in figure 3, where photographs used for correcting small errors in pointing qf the photo- theodolite were not obtained.

Figures 3 and 4 also afford a comparison between the variations of atmospheric pressure and local static pressure l chords ahead of the front airfoil. Except in thc immdiate vicinity of Mach number 1.00, where abrupt changes in local static pressure are quite defi- nitely indicated, the two pressure measurements agree witiin the prhah1e limit of accuracy of the telemetering s y stem.

Because of this limitation on accuracy, further tests must be made before the validity of the smaller differences can he definitely established. The differ- o ences between the subs nic values of atmospheric and local static pressure in figure 3, however, suggest the presence of a blocking effect caused by the airfoils, although the magnitude of this error is larger than would normally be encountered at low subsonic speeds.

CONFIDENTIAL

6 C0WIDENTIAL NACA R1.1 ]To. L6KO8c

The results of the testsare summarized in figure 5, where curves are presented that-show the variations of D/Fp and C D with flight Mach number for the rectan- gular and for the swept-back airfoils. The results for the front and rear airfoils are presented separately because interference effects between body and airfoils and between airfoils may cause the small discrepancies.

Because of the pcssihiltty of interference affects, data for the front airfoils should be the more reliable, par- ticularly at supersonic speeds.

The accuracy of the values of D/Fp shown in figure 5 varies throughout the drop from abOut ±0.02 at M 0.85 to about ±0.006 at M'= 1.20. This variation is due to the increase in atmospheric pressure during the fall of the test bodies and to the fact that the airfoil drag was determined with constant accuracy (±3 lb). Corresponding values for the accuracy of are aboUt ±00035 at 0D M = 0,85 and about ±00005 at M = 1.20. The accuracy with which flight Mach number was determined was about ±0.01, but since the velocity was determined with respect to ground reference the effect of wind has been neglected.

This effect may cause the Mach number determination to be slightly more inaccurate than is indicated by the fore- going value at the lowest Mach number for which results are presented. Such error rapidly becomes negligible with increase in Mach number because the flight path of the test body quickly departs from horizontal during its fall, and, in general, wind velocities are less at the lower altitudes.

The curves of D/Fp in figure 5 show that for the rec- tangular airfoil the drag per square foot of frontal area increased abruptly from 0.04 of atmospheric pressure at a flight Mach number of 0.85 to 0,42 of atmospheric pres- sure at a Mach number of 1.00 and then increased at a slower rate to 0.61 at a Mach number of 117. For the swept-back airfoil the drag per square foot of frontal area increased from 0.02 of atmospheric pressure at a flight Mach number of 0.85 to . 0.26 of atmospheric pressure at a Mach number of 125 without evidencing the abrupt drag rise characteristic of the rectangular airfoil. The drag per square foot of frontal area for the swept-back airfoil was less than 0,15 that for the rectangular airfoil at •a Mach nwriberof 1.00 and less than 0.30 that of the rectan- gular airfoil at a Mach number of 1.17. The measured difference in total'drag of the two bodies agrees with the CONFIDENTIAL

NACA RM Po. r6:Ko8c.

C0FIDENTIAL, measured difference in airfoil drag' within the limits of accuracy of the accelerometer (±0.09 in D/Fp it a Mach number of 0.85 and ±0.025 in D/Fp at a Mach number of 1.20).

Figure 6 shows the effect of aspect ratio on the drag characteristics of rectangular and swept-back air- foils. The variations of D/Fp with Mach number for the p resent rectangular airfoils and for the rectangular airfoils of reference 2 are given in figure 6(a). These curves show that increase of the nominal aspect ratio from 5.1 to 7.6 reduced the Mach number at which the drag rise started by ab o ut 0.02 and' increased the values of D/Fp above the drag rise, by about 0.03. This effect of aspect ratio verifies the results reference 3, which of indicate that a delay'in the drag rise and a reduction in drag at supercritical speeds can be realized for rectangular wings through decrease in as peôt ratio.

Similar variations of D/Fp with Mach number for the present swept-back airfoils and for the swept-back airfoils of reference 2 are given in figure 6(b). These curves show that increase in aspect ratio of the swept- back airfoils from 3.6 to 5.4 resulted in a decrease in D/Fp in the investigated speed range.. This condition may indicate flow distirbancesat the root or . tip of the airfoils where the ideal three-dimensional flow conditions around a swept-back airfoil of infinite span do not exist.

The drag resulting from disturbances at the root or tip would be a greater part of the total drag for the low- aspect-ratio airfoil. It may be noted from the curves in figure 6(b), however, that the values of D/F'p for the high-aspect-ratio airfoils show a tendency to approach those of the low-aspect-ratio airfoils at the highest Mach number investigated.

COaCLUDING REMARKS Drag measurements at transonic speeds attained in free fall from high altitude have been made on rectangular airfoils and on airfoils swept back 45 0 . These airfoils had NACA 65-00 sections, a 60-inch span, and an 8-inch chord rendicular to the leading edge. They were mounted on a 10-inch- CONFIDENTIAL

8 CONFIDENTIAL NACA RM No. L6K08c

diameter cylindrical body. The results of the tests showed that the drag er square foot of frontal area for the swept-back airfoil was less than 015 that for the

rectangular airfoil at a Mach rnmher of 1 O 0 and less

than 060 ;hat of the i-ectangular airfoil at a Mach number of 117 A comparison of these results with the reEults of the previous tcsts on 14O-inch-span NACA 65-009 airfoils having 8-inch chord shows that

1..5 0 increase in

(1) For the airfoils swept back

to 5.4 produced an appreciable

aspect ratio from 3.6

1.2 reduction in drag between Mach numbers of 0.95 and but only a slight reduction at the highest Mach number reached in the tests.

(1,25)

(2) For the rectangular airfoils an increase in reduced the Mach number

aspect ratio from 531 to 7.6

at which the drariss started by about 0,02 and resulted in somewhat higher drag throughout the speed range investt- gated Langley Memorial Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va.

REFERENCES Bailey, F. J., Jr., Mathews', Charles, W., and Thompson, 1.

Jim Rogers: Drag Measurements at Transonic Speeds L5E03, on a Freely Falling Body NACA ACH No. 1945.

Jim Rogers: Com- 2. Mathews, Charles W,, and Thom p son, parative Drag Measurements at Transonic Speeds of Airfoils NACA Rectangular and Svept-Back

650O9

Mounted on a Freely Falling Body. NACA ACR No. L5G30, 1945.

Stack, John, and Lindsey, W F.: Characteristics of 3.

Low-Aspect-Ratio 'Wings at Supercritical Mach Numbers, NACA ACR No. L5J16, 19450 CONFIDENTIAL

NACA RM No. L6KO8c Fig. la,b CONFIDENTIAL

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NACA LMAL.

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(a) Rectangular airfoil.

(b) Swept-back airfoil.

Figure 1.- General views of airfoil test bodies.

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Fig. 2

NACA RM No. L6KO8c

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NACA RM No. L6KO8c

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Figure 3.- Time history of free fail of 1322-p .ound test body

mounting rectangular airfoils. NACA 65-009 airfoil sec-

tion.

NACA RM No. L6KO8c

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• F91 I I I I Mach number, M NATIONAL ADVISORY CONFIDENTIAL COMMITTEE FOR AERONAUTICS Figure 5.- Variation of airfoil drag coefficients and D/Fp with Mach number for rectangular airfoils and airfoils having 45 0 sweepback.

NACA RM No. L6KO8c Fig. 6a,b Restriction/Classification Cancelled

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NATIONAL ADVISORY Restriction/Classification Cancelled COMMITTEE FOR AERONAUTICS.

Figure 6.- Effects of aspect ratio on variation of D/Fp with Mach number for rectangular airfoils and airfoils having 450 sweepback'.