RESEARCH MEMORANDUM F’RELIMINARY INVESTIGATION OF THE TRANSONIC CHARACTERISTICS

OF AN NACA SUBMERGtiD INLET

By John A. Axelson and Robert A. Taylor

Ames Aeronautical Laboratory Moffett Field, Calif. .

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

WASHINGTON June 5, 1950

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NACA RM A5OCl3 __ -.

# NATIONAL ADKEORY COMMITTEE FOR AEROlDWTICS

PRELIMIHARY RWEsTIGATION OF TEE TRANSONIC CEARACIERISTICS

By John A. Axelson and Robert A. Taylor

A preliminary inveetig%tion of m.NACA ~ubmerggd inlet operating over a range of ma884low ratfoe and ancomfng flow angle8 was conducted throu& a Mach number ran- from 0.70 to 1.15 by the u8e of a traneonic b-0 Ram recovery and pressure dietribution were measured for maes- flow ratios up to 0.67. For approximately constant mass-flow ratio, the ranr-recovery ratio decreased about 0.05 in the Mach number ran@ fram 0.85 to 1.1, but generally improved above a Mach number of 1.0 or 1.1. The ram+recovery ratio decreased about 0.05 when the angle tetween, the inlet center plane and the free stream was increased fram 0 to 4 , but fncreased about 0.02 from thie reduced value when the angle wa8 increased from 4’ to 8’. &c??eaeIng the maea flow into the inlet increased the ram recovery, but the improvement became progreesively lees at the higber.ma88 flow8 and higher Mach numbers. Static-preeeure andtotal-pressure 8urveys inside the inlet indicated that the loeees inramrecoverywere caused

-principally by the entrance of low-energg air from the surrounding bound- ary layer which paseed over the aharp ed@s of the-rap wall8 and mixed with the higher-energy air entering the inlet.

IXTRODUCTION

. The locatim of air lnlete on the sides of the fuselag of jet-

PrOpelbd aircraft ha8 received 8peCfa1 emphaEli8 recently because Of the neceseity of housing radar and armament fn the fuselage noflee, Although a side location generally titroducee boundary-layer probleme, a dietinct advantae is sined by the ehorter internal ducting from the air inlet to the compreseor. A8 a result of wind-tunnel iie8ts directed toward the development of a' eide inlet having high pre8sure-recovery characteristics and rnlnjmum adverse effect8 frclm the fu8ela@ boundezy layer, the HACA 8ubmer@d tilet wae conceived. Several variation8 of this inlet were tivestiegted inone of the Ame8 7-by104ootwindtunnel8andare dis- cussed in reference 1. A de8ign judged from the result8 of those teets tObe OpttiUBl~8theIlteSted oTLaWfng-bOdyCo?IIbination in the AU.Bfl

2 a- ITAC!A FM A5OC13 P

164oot high-peed wind txuuzelup to a fr&e+treem Mach number of 0.875. This free-&ream Mach number comeeprmded to elightly higher localI&ch numbers at the inlet, depending upo~l the location of the inlet with respect to the wingandfuselage. (Se 8 reference .2.) The preltiInary lnvestlgs- tionreportedhereinuas conductedusinganidenticaltiletmountedonan almoet flat, two4imeneional surface of a transonic bump in the Ames l& foot high+peed wind tumml, For these test8 the local Mch number over the bump in the re@on of the inlet ran@d from 0.70 to 1.15.

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M

m

P

P

9

u

U

ii2 Al

The symbols used in thie report and their definitions axe as foILlowe:

crose-eectional area of duct, equare Inches

Inlet depth, 1.6 fnches

total pre88ure, pounds per equare foot

boundary-layer parameter deei@ating the height for which a

complete loss of QlWnic pY%SSUre ( >

gJ0’ would be equivalent

to the integratedloss of totalpremure in the actual boundary

Mach number

ma~w flow (flu), aluga per aecand

PXWSSUre coefficient

. . -..

1 .

static preesure, pound per square foot .

dyLWliC pZW88Ure k&J2 ( >

, pounds per equare foot

velocity outeide boundary layer, feet per second

local velocity, feet per second

increment of boundary--layer thicImes8, inches

ratio of duct crose--eectional area &2 inches down&ream of lip leading edge to croee-sectional $rea at rak

"+A RM A50C13 *- 3 ‘L

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%--PO H, -po

rem.covery ratio

ml mo

ratio of the ma88 flowtJIroughthe inlet to the ma88 flow in the

free etream through an area equal to the Inlet area

a angle between the Inlet center plane and the free stream (8imuhting the an& Of attack Of an airplane Side inlet>, degreee

boundary-layer thiChe88 Where the local Velocity i8 0.99 Of the velocity outeide the boundary layer, tiche

s* boundary-Layer di8plaCement thiCkle88

boundary-layer momentum thiCkne88 [ l"z (1-e) ti]s incha

ma88 density, 81x.1@ per cubic foot

Subscripts

averagg condition8 over test sectian of bmp

diffuser entrance

duct rake

APPARATUS

Deecription of Tnlet Mode1B

Detail8 and dimensions of the NACA submerged-inlet model are ahown in figure 1. %80far 88 possible, the dimension8 of the inlet COrY'e- sponded to those of the inlet reported in reference 2, where the inlet wa8 Installed on the curved side of a modelfusela~. 32 the pre8ent investigation, the inlet wag mounted on a Mo-dimensianal surface'ae ahown in figure 2. To siu@ate angle-of-attack conditicmrs of an aIrplane

NACA RM A50C13'

side inlet, the inlet was mounted on the transonic bump with angles of O", 4O, and 8’ between the Inlet center plane and the f me stream. The cur- vature of the test surface necessitated the construction and installation. of three separate models with identical basic inlet lines.

The models were equipped with pressure orifices alang the center line of the ramp andaround the lip (exzeptthe inlet representing8'angle of attack which had no lip orifices), Pressure losses and flow rates In the Inlet were measured with a rake 6 inches behind. the lip leading edge. !The rake consist&d of 30 total+ressure and 35~stat~c-press-ure tubes. The a* which flowed through the inlet entered a diffuser which started 6 jnohes downstream of the rake and discharged back'to the wind-tunnel air stream through the underside of the bump.

Description of the Transonic Bums

The trsnsonic bump of the Ames l64oot high-speed wind tunnel used for testing models through sonic velocity and up to low supersonic speeds is sham in figure 3. The bump had an l&foot chord and a flat tmderside which was mounted a small distance away from one of the vertical walls of the wind tunnel. The profile of the bump was essentially one-half of an NACA 16421 section modefed by a 17-percent extension of the chord and faired by a straight line connecting the 64-percent-chard point of the resulting profile to the trailing edge.

Distributims of local Mach number over the bump surface are shown in figure 4. At the highest Mach numbers, there was an increase in the local Mach number and consequently a small favorable pressure gradient alongthatportion of the bump surface inwhich the submered Inlet was . placed. The maepitude bf the favorable pt%ssure gradient on the blimp was, however, small compared to the gradient of pressure along the ramp of the fnlet, amounting.to less than 3 percent of the gradient on the ramp below a Mach number of 1.05, and less than 7 percent at the highest Mach numbers, Thus, the gradient of Mach number was felt to have only a small effect on the results obtaIned for the Inlet, and was of a maep nitude which could conceivably exist along the side of the fuselage of an airplane.

The underside of the bump was mounted a small distance away from the vertical wall of the wind tunnel in order that the boundary layer of the t~lwallwouldpassundero the bum@, A 2--inch spacing existed during the &ests 05 the inlet at 0 angle of attack, When the inlet was tested at 4 and 8 angles of attack, the spacing wae increased to 5 inches in order to reduce the boundary layer on the bump and to decrease the static pressure under the bump so that slightly hi&er mass flows through the inlet might be obtained. (The duct exhausted through the

underside of the bmp.) Results of boundary-layer survey8 7 inches

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NACA RM A5OC!l3 5

h forward of the rRmp at bump station 73.3 are presented in figure 5 for the bump located2 inchee and5 inchesfromthe wind-tunnelwall. The buq boundary layer was decreased by increasing the spacing, but time did not permit repeating the tests of the inlet at O" angle of attack with the 5-inch spacing.

TESTS

A Mach nmber rang8 fram 0.70 to 1.15 kae covered in the investi@- tion of the inlets, the Msch numbers being taken as the average of the values prevailingbetweenbump stations 78 and 114 inches. Vnder the test conditions, this ran@ of Wch numbers corresponded to a Reynolds number range from 3.5 to 4.2 million per foot of length.

Three different mass--flow conditions were investigated with the inlet at 0' angle of attack. The two reduced flow rat88 were produced by the additirm of constrictions 13 inches behind the lip in the diffuser entrance. Since rigid control of the mass flow during the tests was not practicable, there were small wriatians in the r8sultlngmass~low ratios over the Mach number ran@. Only one mass4low condition was investimted at 4O and 8' angles of attack.

RESISLTS AND DISCDSSION

Ran&Recovery Ratio

The ram--recovery ratios were computed by the method outlined in ref- erence 2 wherein the logarithms of the local total pressures at each of the 30 total-ressure tubes of the rake were weighted according to the local mass flows.

The primary variables which affect the ram recovery of the inlet and which can be isolated in the present investigation are maasdlow ratio, angle of attack, boundary-layer profile, and Mach number. Th8 effects of each of theee on the ram recovery will be discussed,

Effect of mass-flow ratio.7 The variations of ram-recovery and mass- flow ratios with Mach number for three different diffUse3X3ntranCe con- stricti& are shown in figure 6 for the inlet at an angle of attack of 0'. A cross plot of these results is presented in figure 7. At all test Mach numbers increasing the mass4low ratio resulted in an increased ram recov- ery, but the improvement generally became progressively mailer at the higher mass-flow ratios and at the higher Mach nmbera. These results are in agreement with those measured during the investigation reported in

6 NACA RM A50C13 .-

reference 2 wherein the optimum mass-flow ratio was about 0.70 over the Mach number range from 0.30 to 0.875. S&e only one flowconditiaawas tested with the inlet at angles of attack'of 4' and 8’, it was not possible to present the rmecovery ratios for these angles of attack at constant values of mase-flow ratio. The amount by which the mase-flow ratios varied, however, were not lar@ enough to obscure the consistent trend In the var- iation of ram+recovery ratio with Mach number.

Effeot of angle of attack.- The variations with Mach number of ram- recovery ratio and mass+Plow ratio for angles of attack of O", 4', and 8' ' are shown in figure 8. Because of th8 dffference in spacing b8tW88n the underside of the bump end the wind-tunnelwall, slightly higher mass flows were obtalaed with the Inlet at 4' and 8' angles of attack than were obtained with the inlet at 0' angle of attack. The mass-flow ratios shown in figure 8(b) are not to be compared forevaluation of an effect of angle of attack, but are shown in order that a comparison can be made between the ram-recovery ratios shown in figure 8(a). Had the same bump Spacing b88n used for all three angles of attack the mass flows would in all probability have been nearly equal at any given Mach number. The small differences in measured mass flows between the 4' and 8' angles of attack w8re probably within the accuracy of the test results.

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The results show, at least qualitatively, that the ram+8 oovery ratios for the 4' and 8’ angles of attack were slightly less than those for 0' angle of attack in spite of the fact that the masaiflow ratios were higher by 0.02 to 0.09. A c~~~parison probably more quantitative in nature is possible by performing an extrapolation of the results shown in figure 6 for th8 inlet at 0' angle of attack so that the ram-recovery ratios for three angles of attack might be compared on the baais of equalmass4low ratio. (The results presented in figure 8 of reference 2 which cover higher mase-flow ratios than those covered In the present investigation indicate that serious error is unlikely in making such an extrapolation.) For example, in the present investigation at a &iach number of 0.95 the ram+ecovery ratio for 0' angl8 of attack extrapolated up to a masa+i?low ratio of 0.62, the value shown in figure 8(b) for 4' and 8' angles of attack corresponds to a value of about 0.83; but introducing this value in figure 8(a) has little effect on the relative values of the ram+e c every ratios for the three angles of attack. At ccmmrabls mass-flow ratios, the ram+recovery ratio decreased about 0.05 when the angle of attack was increased from 0' to 4', but -roved-about 0.02 when the angle of attack was increased frcxn 4' to 8’. A similar variation of ram+c%covery ratio with angle of attack was reported in figure 9 of reference 2 for the inlet fitted with amall boundary-layer deflectors shown in figure 3 of refer- ence 2. . --

Effect of Mach number and bump boundary layer.- The determination of ,the ram-recovery characteristics of an RACA i3'ubm8r@d inlet through the transonic speed rang8 tis the primary purpose of the present investigation and the reeults have been presented in fig&es 6(a), 7, and 8(a). The

3

NACA FM A5OC13 7

-- effects upon rsm+ecovery ratio of variations of the mass4l.o~ ratio and of the angle of attack have been disoussed, leaving the effects of Mach nuder to be isolated. Since the results presented in figure 5 indicate thatchan@s occurred inthe boundarylayer onthe bumpwhenthe Mach number was varied, it is desirable to ascertain what effect the boundary- layer chan@s'exerted onthe ranr-re covery characteristics of the inlet.

As shown infigure 5(d),the maximum&an@ in the boundary-layer parameter h/d throughout the test range of Mach numbers was about 0.025, most of which occurred below 0.92 Mach nmber. The results fram the pres- ent investigation have been cabined in figure 9 with those fra refer- ence 2 and show the relationship between the boundary-layer parameter h/d and the r Wcovery ratio of the inlet. The results for the three Mach numbers for which'the two investigstione overlapped indicate that the 0,025 &an@ in h/d could cause an increment of rwecovery ratio of about 0.02. However, in th8 present investi*tion, the boundary-layer parameter h/d remained almost constant above 0.92 Mach number, while the most significant changes in ram-recovery ratio occurred above this Mach number. It appears, then, that in enalyzing the variation of ram+oe covery ratio with Mach number for Mach numbers above 0.92 it is permissible to consider the bump boundary layer essentially constant, thereby allowing further scrutiny of the effects of Mach number.

The results in figure 9, in addition to providing evidence on tbs effect of the changes Ia the boundary layer, also serve to correlate the ram+recovery results of reference 2 with those of the present investiga- tion. Perhaps more important, however, the results in figure 9 showhow large an effect the energy deficiency in the oncoming boundary layer exerted an the ram+recovery characteristics of the inlet.

It has been shown that the effects of changes in the mass4low ratio and in the bmp boundary layer above 0.92 Mach number were too amall to mask the consistent reductions in ram-recovery ratio which oc&rred at Mach numbers near 1.0 and the slight improvement above l.lMach number. Zif'ormation on the changes in the flow into the inlet which accompanied the changes in the Machnmberis introduced in the followin g sections, which present the distributicne of the losees in the inlet, photographs of tufts on the model, and pressure distributions.

Distribution of rsecovery losses inside the inlet.- ti order to show the distribution of the r Mcovery losses within the inlet, con- tours of the cmput8d local r-ecovery and niass-%low ratios are pre- sented in figures 10, ll, and 12. The results are arranged to show pri- marily the effects of Mach number, the mass4low ratios for each set of three Mach numbers being chosen as nearly equal as possible from the available data.

The results shown in figure 10, the intepated values of which appeared in figure 6, cqver three maes4low conditions for O" angle of

8 NACA RM A5CC13

attack, As would be expected, the regions of high ram recovery corre- sponded to those of high mass flow. B gmeral, the losses were con- centrated in those-areas adjacent to the ramp and in the corners where boundary layers were to be expected. The asymmetry of.flow and of losses which were mor8 evident at the lowest mass flow (fig. lO( g)) was probably caused by small differences between the boundary layers in the upper and lower portions of the inlet rather then by a change in the flow direction over the inlet, dnaamueh as the latter would have had a similar effect on the results at the higher maas flows and, further, tuft studies indicated no change In flow directian over the bump.

Theram-re and8"

covery andmass4low contours for angles of attack of 4' are shown fnfigures lland 12. Asmightbe ewected, there was

a concentration of the losses in the lower portion of the inlet because of the differences In the direction and spilla@p3 of air and in the bound- ary layers along the two diverging ramp walls. Differences in the bound- ary layers would be expected in light of the differences ip the pressure distributions presented in fim 13, which were measured during the investigaticm reported ti reference 2. These results show l-f&r@ diffeYWIC88 between the local pressures along the two:walls when the inlet was opera- ting at other than zero angle of attack.

Cause of the lOSS8S.- In reference 3J: it was sumised from re recovery distributions for a similar inlet that the flow of air over the sharp edges of the ramp walls and into the inlet Imparted a rotational velocity component to the air. It was reasoned that this rotational velocity increased tith maes-flow ratio end with the divergence of the ramp walls relative to the direction of the air streamand resulted in the formation of one or more vortices, the centers of which produced regions of low local ram recovery not immediately adjacent to the walls. In th8 present investigation, similar regions of ram+ecovery loseee were measured, such as those shown in figures 11(a) and X?(a). In order to provide more information on the nature of the ram+8 covery losses, the contours of figure ll(a) are compared with the measured total and static pressures across the inlet ti figure 14. In light of the results shown in figure 9 and the large variations In total pressure across the inlet, it appears logicalto attribute the regions of ram+x covery loss and low total pressure to the deficiency of energy in the air which cam8 from the bump boundary layer and left the surface in passing over the sharp edges of the ramp wall before reaching the inlet. O&Ly relatively small Par- iations of static pr8ssure across the Inlet and no marked reductions in local static pressuP Such as might be expected at the core of a vortex were measured in the present investigation, so it is possible that the rotational velocity ccrmpanents remained as more or less random vortiofty or turbulence. It Is also possible that, at the massSlow ratios covered in the present investigation, the cores of.the vortices passed outside the Inlet. The latter would explain oondensation trails which were observed In the wind tunnel during the course of the tests. With higher maas-410~ ratios than were obtained in the:present investigation, it is

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NACA B&f A5CC13 9 -*

-- possible that the rotational velocity components could wrap up into vortices which would enter the inlet. Cn the basis of the present inves- tigation, however, it is felt that the losses in ram recovery were, in the most part, caused by the entry of air from the bump boundary layer.

The diverggnt rams of the HACA~submerg8d inlet appears to offer two advsntages with respect to the boundary layer in which it is placed. First, the oncoming boundary layer that flows directly on to the remp does not thicken, aa it approaches the inlet, in COrr8SpOdenCe With the adV8rS8 pressure sadlent it encountirs becauee the diverging walls provide relief in a third dimension; and, second, the boundary layer that flows frm the outside surface on to the ramp by passing over the Sharp edges of the ramp wall is involved in a mixing action with higheMnergy air and SO IS less prone to separate.

Pressure Distributions and Tuft Studies

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In the preceding discussian, the ram-recovery lossee w8re attributed principally to the entrance of low-nergy air fram the bump boundary layer. The bump boundary-layer parameters varied comparatively little above 0.92 Mach number; however, the ram-recovery ratios varied in the Mach number rang3 fram 0.92 to 1.15. Tuft studies, scpII8 photographs of which are pre- Sented in figure 15, indicated that no si&ficant separation occurred on the ramp. It appears logical, then, to attribute the variations in r8I11- recovery ratio with MachnMber primarily to chsnges in the amounts of low-energgairwhichentered the inletfrcm the surroundingbump boundary layer. The controlling factor which determined the amount of bump boundary- layer air which entered the inlet was the local pressure gradient between the ramp and the Surroundingbtmrp surface. Pressure distributions along . the ramp, .some of which appear in figurea 16 and 17, indicate that the pressure differenceS were generally greatest and extended over longer portions of the ramp at the Mach numbers corresponding to those at which the minimum ram+ecovery ratios were measured. (See figs.7and8.)

The effect of variation in mass+Ylow ratio m the pressure distribu- tions along the centerline of the ramp andaroundthe lip is abo shown in figure 16 for an angle of attack of Co at Mach numbers of approximately 0.75, 1.02, end 1.14. Reduction of maes flow at the higher Mach numbers bad the eqpected effects of increasing the pr8SSLUXS on the ramp and increasing the angle of attack at which the lip operated. Figure 17 pre- sents additianal pressure of O", 4', and 8O

distributionrs along the ramp for angles of attack and around the lip for 0' and 4O. Varying the angle of

attack with a constant mass flow had a noticeable effect on the pressures around the lip, but had little effect on the pressures along the center line of the ramp. It Should be repeated, however, that variation of angle of attack produced large changes in the pressures along the ramp walls as shown in figure 13.

10 . NACA RM A50Cl3

Since the pressure distributions and tuft studies indioated that the entrance of air from the oncoming boundary layer occurred principally over the forward part of the ramp walls where the local pressures were less than those of the surrounding air stream, it appears that, if boundary- layer control were to be employed, the surrounding surface just outside and forward of the ramp should be COnSid8r8d in addition to the surface of the ramp itself. The tufts shown in figures 15(a) and 15(b) indicate that the boundary lapr along the after part of the rm walls passed outside of the lip in the direction of the local pressure gradient.

CCNCT.DSIONS

The following conclusions were drawn from the results of tests of an NACA submerwd inlet on a trsnaonic bump for a test Msch nuuiber range fram 0.70 to 1.15:

1. For mass+Plow ratios between 0.40 and 0.67, the ram-recovery ratio decreased about 0.05 in the Mach number range from 0.85 to rou&ly 1.1. Generally there was a small improvement in the ram-recovery ratio at Mach numbers above 1.0 or 1.1.

2. For comparable mass-flow ratios, the ram+recovery ratio decreased about 0.05 when the an& of attack was increased fram O" to 4', but improved about 0.02 when the angle of attack was inoreased from 4' to 8’.

3. Increasing the mass-flow ratio resulted in higher rsm+recovery ratios but, in general, the improvement became progressively less at the higher mass flows and Mach numbers investigated.

4. ,Rem+recovery ratios higher than those obtained in the present investigation appear possible on installations with relatively thinner boundary layers.

Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics,

Moffett Field, Calif.

t NACA RM A5OCl.3

1. Mossman,lQrmetA., andRmdaU,LaurosM.: AnEqerhentallh~sti~- ti'on of the Desiep Variables for NACA Submerged Duct Entrances, HACA RM A7I30, 1948.

2. H&U., CharlesF., andBarclay,F.Dorn: An Xkperimental Investigation of NACA Submer@d Ihlets at El& Subsonic Speeds. I - &lets Forward of the Wing Ieading Edge. HACA RM M~16, 1948.

3. &hny, NO81 K. : An Investigation of Submerged Air Inlets on a l/h- &ale Model of a mica1 Fighter4ype Airplane. NACA RM A8A20, 1948.

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NACA RM A5OCl3

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Section A-A

Secfion B-B Lip Coordinofes

Figure f.- Dhensiono/ dafu for mode/ h/ef. v

Afofe: A/i dimensions in inches

Romp-Wail Coordinates

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(a) wpstreanl view. (b) Dovnstream view.

Figure 3.- Moael of NACA Buhmerga tit molmted on the tmmonlc bw in the Aslee l64oot hl&+qeea wird tunnel.

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Bump chord ) percent 45 50 55

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92 96 IO0 IO4 108 l/2 116 I20 Bump station, in. v

Figure 4.- Mach number distribution perpendicular to the surface of the tronsonic bump.

/6

Bump chord, percent

45

I’ Y Y

80 84 88 92 96 IO0 104 108 II2 116 120

Bump station; in. v

(bl Me, 0.9Z

Figu r 8 4. - continued .

.’ ,’ ,’ .’ .’ .’

PO

16

Bump chord, percent

40 45 50 55

I I l.d4

x / 1

I I /- \ \.

I / /

I ” I \ I I Y I’ I I v I

I /I I/ t /-- \

II / 1 /I 1 /I

1 I InI I /I I

\ \\

76 80 84 88 .92 96 IO0 104 IO8 II2 II6 I20

BUmQ station, in, v

Id M,, 1.12.

Figure 4. - Concluded.

22 NACA FiMA~Cl3

2” Gieurance -------- 5” Clearonce

.8

.6 (a) Boundary- iuyer thickness.

.08

.06

-08

.06

(b) Displacement Mickness.

fc) Momentum fhickness.

./6

- - -.--.-

19 1 I I t I I I I I . ,L

.68 .76 .84 .92 1.00 LO8 I. f6 Mach number. A4,

(d) Boundary-iayer parameter #. T

figure 5.- Varhfion with Moth number of. fhe bump boundary- layer choracferistics measured at 34-percent chord for two clearances between fhe bump and wind - funne/ ~a//.

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RACA RM A5OCXL3 23

.86

.82

-78 p” Q.. a? &Q I

.74

I i I I I h

(0) Ram-recovery rufio.

30 I I

40

-60

.85 -90 -95 I.00 f. 05 i.iO f. 15 Mach number, M,

(6) Moss -f/o w ratio. -597

Figure 6.- Vurhtion of rum-recovery ratio and mass- f/ow ratio with Much number for f/tree diffuser-in/ef areu rofios. a, Of

.70 .80 .90 I.00 1.20

Mach number, M,

Figure Z- Voriathn of ram-recovery ratio wifti rWach number for ssveraf mass-flow ratios. tz, 0”

NACA RM A5OC13 25

.86

(0) Ram -recovery ratio.

.70

.60.

-- -30 .- - 70 -75 -80 .8.5 20 -95 LOO i-05 /. 10 /./5

Much number, M, -

(b) Mass-f/ow r&o.

Figure 8.- Vurhtion of rum- recovery rUti0 and muss- flow fotio wifb ~ffch

number for ungles of olfuck of ,04, 49 und 8*.

26 NACA BM A5OC13 .-

t? 2 0 .88 5 I

e Q A?4

.80

o 0.80 0.58

I I I I t 1

! 0 -05 .I0 .I5 .20

Bumdory - foyef pofomefet $

:

Hgutt? 9 . = Voduf/on of tom- recovery tdo with boundaty- layer patam eter h

a* a s 0 S

RICA RM A50C13

2 s

(a) M,, 0.70 ;

Lzlr m. > 0.57.

Ram-fecovery f afio

Mass-f/o w f ofio

fi 1 M,, i-02; W% K-g # 0.80;

ml m* ’ 0.54.

Figure lo.- Ram-recovery and moss-flow ConfOUB in the e??fYat?Ce of the submerged in/et, a, 0” v

27

28 ITACA RM A5OC13

fd) 4, 0.72; Y-4 J%Yg

I 0.81;

2 m. f 0.40.

ffum-recovery ratio

Muss-f/o w I utio

(e) MO, i-O/; B ~- 4-e

I 0.77;

2 m, jO.40. .

Figure IO.- Continued,

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NACA RI4 A50213

Rum-recovery rafio

Mass-flow f of/‘0

(9/ N, 0.76; Y-e

H-p a 0.70; 0

& m. # 0.25.

29

(i/ 4, /./4; 6-P H-p t 0.76; * 0

2 ’ o-30-

Figure /O.- Concluded.

30

Ram-recovery fafio

Y-J? - # 0.79; Ho-4

m &* > 0.66.

Moss-flow fofio

6, f 4, 0.89; t%

H,-% I 0.80;

# I 0.63. 4

NACA RM A5OC13 .4

figure //.- Rum- recovery and muss- flow canfoufs in the en - ffonce of fhe submerged Mef. CI, 44 v

XAC!A RM A5OCl3

9 kg

’ 0.83;

Ram- recovery f afio

Mass-ffow rafio

31

Figure /2.- Ram- recovery and mass-flow confours in the enfrance of fhe submerged infef. CT, 83 v

-.8

.8 10 .20 30 40 IO 20’30 ,4u

Fuselage station, @.

I&) Upper wolLi

0 0” 0.55 A 0” .90

El 4O .55

0 4” .90

fc) Lower we/l:

Figure /3.- PfeSSum distfiitruon along ramp and diverging hmp wa//s fm jnves~@jon reported in

reference 2. K, 0.80.

-,

l -

NACA RM A50Cl3

Ram-recovery ratio

4% - 2 0.83; - H,-p, 2 I 0.65. 0

Mass-flow rofio

31

(k) It& f-13; Y-42

- s 0.79; Ho-4

2 s 0.56.

Figure @.- Ram- recovery ond mass-flow confours in the entrance of the submerged’ inlef. CT, 89 v

w

-.8

.8

Ial Romp.

Fusefoge station, in.

fcl Lower no//.

Q 0” 0.55 A 0” .90

q 4” .55

0 4” .90

Figure i3.- Pressure distribution t&g ramp and diverging ramp wa/ls from investigation reported in

reference 2. I%, 0.80.

: ‘. ‘I

NACA RM A5OC!l3 . 33 -_

l -

.

l -

--

6

5

4

3

2

/

0 I I I k m I I L I I

24 26 28 30

Local pf essure , in. Hg.

Figure /4.- Compurkun of /ocal pressures and fcwn -

I8 CO Vet-Y TU f/OS Wf=fhh fh8 id8f. M, t cf. 71; Q, 4 = ;

HI-&,,o.79; 2, 0.66. ffo-& 0

.

.

NACA EM A5OCl3 35

(a) MO, 0.78; 2, 0.58.

t

(c) MO, 1.05; 2 , o-53.

Figwe 157 Tufts on the submer~d-Inlet model at several' Mach numbers. a, 0'.

NACA RM A5OCl3 37

. (a) N,, 1.13; 2, 0.53.

Figure 15.- Conoluded.

P r I’ I’ .

-.8

-.6 0 0.74 0.85 0.58

4 8 t2 I6 20 Ramp station, in.

(0) Ramp pressure distribution.

24 0 .5 I.0 1.6 Ll;o station, In.

0)) L@ pressure distribuiion.

Figure /6.- Ramp and lip pressure distributions for several muss- flow rafios

and Mach numbers ; CT, 09

-1.0

-.8

-.6

2 .4 2 a .6

’ .8

/.O

1.2

‘. ‘.

0 4 8 12 16 20 24

Ramp station, in.

fcl Ramp pressure dkti$~ion.

Figure /6,- Continued.

1 ‘I

0 .5 I.0 f.5

Lip station, in.

Id/ Lip pmssure distribution.

4.0

-.8

-.6

-,4

-.2

0

.2

.4

.6

.8

l-0

f.2

I.4

0 i.15 0.8; 0.56 El I.14 .77 .40- A f,f4 .76 .30

I I I I

L/j2 leading - sdgs

0 4 8 I2 I6 20 24 Romp stothn , in.

(e) Romp pressure distribution.

figure 16. - Conchded.

“

0 .5 I.0 1.5 Lip sfofion, in.

(f) Lip pressure distribution.

4.0

-.8

-.6

-.4

-.2

0

.2

.4

.6

.8

f.0

I.2

I.4

1 Te m, MO H*-& a a 0 0.74 0.85 0.58 0" EI .75 .82 .67 4" A .73 a83 .65

I I I I I 8'.

0 4 8 12 16 20 24 Ramp station, in.

la) Ramp pressure distibuthn.

0 .5 I.0 1.'5 Lip station, in.

(b/ Lip pressure distribution.

Figure 17. - Ramp md lip pressure d/striButions for three angles of attack

oi various Moth numbers.

l. ‘. ‘.

-1.0

-3

-.6

1.0

I.2

I.4

I 0 I.02 0.80 0.54 0" El I.05 .75 .56 4" Al.04 -77 .57 8"

I I I I I I I I I

0 4 8 /2 /6 20 24 Romp siation , in.

lc) Ramp pressure disMbution. ’

Flgure 17: - Continued.

Inner

P 0 .5 I.0 I.5 L/p statian, in.

td/ Lip pressure distribution.

-I. 0

-.8

-.6

f.4

0 I.15 0.81 0.56 0’

0 4 8 f2 I6 -20 24 Ramp statian, in.

(e) Ramp pressure dishbutihn . v

Figure /Z- Concluded.

l . ‘, ‘. ‘.