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SECURITY INFORMATION

RESEARCH MEMORANDUM

FORCE AND PRESSURE RECOVERY CHARACTERISTICS AT

SUPERSONIC SPEEDS OF A CONICAL SPIKE INLET WITH

A BYPASS DISCHARGING FROM THE TOP OR BOTTOM

OF THE DIFFUSER IN AN AXIAL DIRECTION

By J. L. Allen and Andrew Beke

Lewis Flight Propulsion,LaboratoryCleveland, Ohio

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NATIONAL ADVISORY COMMITTEEFOR AERONAUTICS

WASHINGTONMarch 23, 1953

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.-I?ACARM E53A29 ‘

NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS

RESEARCH 141M)RANDUM

FORCE AND PRESSURE RECOVERY CHARACTERISTICS AT SUPERSONIC

SPEEDS OF A CONICAL SPIKE INLE?IWITH A BYPASS

DISCHARGING FROM TEE TOP OR BOTT!OMOF THE

DIFFUSER IN AN AXIAL DIRECTION

By J. L. Allen and Andrew Beke

& axially symmetric naceLle-type conical spike inlet with a fixed---area bypass located in the top or bottom of the diffuser was investigatedin the Lewis 8- by 6-foot supersonic tunnel. The bypass was sized todischarge in a nearly axial direction about 10 percent of the maximummass flow captured by the inlet. Force and pressure recovery data wereobtained at flight Mach nunibersof 1.6, 1.8, and 2.0 over a range ofangles of attack from 0° to 9°.

Top or bottom location of the bypass within the diffuser did nothave significant effects on diffuser pressure recovery, bypass mass-flowratio, or drag coefficient over the range of angles of attack, flightMach numbers, and stable engine mass-flow ratios investigated. At aflight Mach nu@er of 2.0 and angles of attack from 3° to 9°, a largerstable subcritical operating range was obtained with the bypass on thebottom. Higher lift coefficients and nmre positive pitching momentswere obtained with the bypass on the bottom over the range of angles ofattack and flight Mach nunibersinvestigated.

At zero angle of attack and a flight Mach number of 2.0, about14 percent of the maximum stream tube entering the inlet was bypassedwith a drag increase of only 20 percent of the additive drag that wouldresult for equivalent spillage behind an inlet normal shock. Diffusertotal-pressure recovery was not significantly reduced compared withresults obtained without bypasses.

INTRODUCTION

Previous investigations (refs. 1 and 2) of an axially syu.unetricspike-type nose inlet indicated that discharging Mss flow in excess ofe~ine requirements by

*means of a bypass increased the drag by only a

2 -mmm!g,., NACA RM E53A29

fraction of the additive drag that would result for equivalent normal- Ushock spillage and did not significantly reduce diffuser total-pressurerecovery. The data of reference 2 were obtained with two fixed-areabypasses on opposite sides of the model in a horizontal plane, and the “

total mass flow bypassed was about 20 percent of the free-stream tubeentering the inlet. At angles of attack other than zero, various cir-cumferential locations of the bypass may result in significant variationsin performance because of differences in the external flow field near the Nbypass exit as well as internal flow differences near the bypass entrance. EIn addition, bypass mass flows less than those of reference 2, which would

m

be necessary for a variable mass-flow bypass system, may not result inproportional gains in performance compared with normal-shock spillage.Therefore, in order to extend the results of reference 2, the same inletmodel was investigated with one identical bypass installed in the top orbottom of the diffuser. The investigation was conducted in the NACALewis laboratory 8- by 6-foot supersonic tunnel and the results are pre-sented herein.

SYMBOLS

The following symbols are used in this report:

area

maximum external cross-sectional area

drag coefficient, external drag plus internal anddue to bypassing mass flow, D/q&

lift coefficient,measured lift minus internal lift due to engine

—

.

.

external dm.g

mass flow

%%

pitching-moment coefficient about base of model,total minus internal pitching-nmment due to engine mass flow

-.—

*I

CT.D thrust-minus-drag coefficient, (T - D)/~

D dreg force, external drag PIUS internal and extern~ *ag due ___ —to bypassing

L length of subsonic

z over-all length of

diffuser, 46.9 in.

model, 58.7 in.

NACA RM E53A29 3

M

m

%+0

Q4/w

P

P

(Ps/p~)

~

T

v

x

a

Y

P

Mach nuniber

mass flow

bypass mass-flow

engine mass-flow

total pressure

static pressure

bypass or nozzle

ratio,

ratio,

bmass mass flow

POVOA1

engine mass flow--.pov@l

pressure ratio, surface static pressure with-out bypass (station 33.0) per-total pressure of Jet

dynamic pressure, ypM2/2

thrust, net force in flight direction due to change of momentumof engine mass flow between free stream (station O) and dif-fuser discharge (station 4) including balance base force

velocity

longitudinal station, in.

nominal angle of attack, deg

ratio of specific heats for air

mass density of air

Subscripts:

b bypass

x longitudinal station

o free stream

1 leading edge of cowl

4 diffuser discharge at constant diameter section, station 46.9

4,1 diffuser discharge at constant diameter section (sting out],station 46.9

4 NACA RM E53A29

Pertinent areas:

Am maximum external cross-sectional area, 0.360 sq ft

Al inlet capture area definedby cowl lip (measured),0.155 sq ft

4 flow area at

A4,1 flow axea at

The model, which

diffuser discharge, 0.289 sq ft

diffuser discharge (sting out), 0.338 sq t%

APPAR.KWS AND PROCEDURE

was identical to inlet B of reference 1, con-sisted of a single-conical-shockinlet without internal contraction, anannular subsonic diffuser, and a fixed-area bypass which was identicalto the bypass of reference 2 except for circumferential location (fig. 1).Tip projection of the 250-half-angle cone was selected so that theconical shock would intercept the leading edge of the cowl lip at aflight Mach muiber of 2.0 and provide a mass-flow ratio of unity. Atthis condition the streamline behind the oblique shock was nearly alinedwith the slope of the external portion of the cowl lip. Coordinates ofthe cowl and centerbody are presented in table I and the longitudinalarea variation of the subsonic diffuser is shown in figure 2. The arearatio is expressed as the quotient of the local flow area based on theaverage normal to the annulus surfaces and the maximum flow area at thediffuser discharge (station 46.9). The leading edge of the bypass wasapproximately 6 inlet diameters downstream of the inlet entrance andcorresponded to a position slightly forward of the compressor inlet ofa turbojet engfne or the cOfiUstiOn Cmer of a ram-jet engine.

The bypass insert and the outer body, or shell, formed a convergent-divergent asymmetric nozzle, shown photographically in figure 3 and indetail in figure 4, which was capable of discharging in a nearly axialdirection about 10 percent of the msximum mass flow capturedby theinlet. Theoexternal surface of the bypass was a channel set at an

angle of 3$ relative to the model axis of symmetry and did not protudebeyond the external cylindrical contour of the model.

The model, which was sting-mounted from the tunnel strut, had aninternal three-component strain-gage balance. Balance normal andmoment readings were used in conjunction with a static calibration ofmodel and sting to correct the angles of attack for deflections due toaerodynamic loads. Actual angles of attack were as much as 0.4°greater than the nominal angles; however, all data were reduced for thenominal angles of attack. Differences in actual angles of attackbetween the model with the bypass located on the top or bottom were

w

.

.

—

.

—.

NACA RM E53A29 5

within O.1O. Regions of inlet instability, or pulsing, were determinedfrom time-force histories of axial-force variation andby means of high-speed schlieren motion pictures.

The sum of the mss-flow ratios of the engine and the bypass, basedon the mass flow of a free-stream tube defined by the cowling capturearea,”is the mass-flow ratio of the inlet. Methods of instrumentationand calculation are given in reference 2. The accuracy of the enginemss-flow ratio is approximately 1 percent at zero angle of attack andwithin 2 percent at an angle of attack of 9°.

In order to account for the thrust developed between the plane ofsurvey (station 36.7) and the diffuser discharge (station 46.9), thediffusion between these stations was assumed to be isentropic. Themeasured thrust-minus-drag coefficients correspond to diffusion withthe support sting reroved inasmuch as the force (determinedly measur-ing the static pressure) acting on the base of the strain-gage balanceis, within about 1 percent, equal to that obtained by diffusing isen-tro~ically from srea A4 to A4,1. Accordingly, the diffuser-discharge

Mach nurtibersare based on the mea A4,1. The Remolds n~er~ based

on inlet diameter, varied from 2.10 to 2.19W06.

RESUITS AND DISCUSSION

Presentation of Results

The vsriation of bypass mass-flow ratio, total-pressure recovery,diffuser-discharge Mach nuniber,and coefficients of thrust-minus-drag,drag, lift, and pitching-moment with engine mass-flow ratio are presentedin figures 5 to 8 for the bypass mounted in the top of the diffuser andin figures 9 to 12 for the bottom bwass location. Data obtained atflight Mach nunibersof 1.6, 1.8, and 2.0 are presented in figures 5 and9 for a nominal angle of attack of zero and in figures 6 and 10 for anominal angle of attack of 6° for the inlet with the bypass on the topand bottom, respectively. Data for nominal angles of attack of 3° and9° at a flight Mach nwrber of 2.0 are presented in figures 7 and 11,and lift and pitching-moment coefficients for all flight ~ch nmibersand angles of attack investigated are presented in figures 8 and 12.Schlieren photographs showing the flow field in the region of the bypassdischarge are presented in figure 13 for the two bypass locations andangles of attack of 0° and 9°.

The thrust-minus-drag coefficients were obtained from the strain-gage balance readings and correspond to the net force on the model inthe flight direction with sting removed and can be used for general com-parisons of the data. Since the over-all thrust of the propulsive unit

6 NACA RM E33A29

●

is composed of the net forces of the inlet diffuser, engine, and exhaustnozzle, the thrust-minus-drag coefficient can be used directly in com-puting propulsive unit performance. Drag force was obtained by subtract-ing the measured thrust-minus-drag from the thrust computed from the massflow consumedly the engine (see SYMEX&3). The drag coefficient thusincludes the external drag of the nmdel plus the net internal and exter-nal effect due to bypassing mass flow. Similarly, the lift and pitching-moment coefficients are the difference between the measured value and thecomputed internal lift or pitcm~ moment caused by the er@ne mass flow.The additive components due to mass-flow spillage behind the inlet shocksystem are included in the drag, lift, and pitching-moment coefficients.Pitching-moment coefficientswere computed by assuming that the turningof the engine mass flow occurred at the cowl lip.

Effect of Top or Bottom lbcation of Bypass

For symmetricalbodies at positive angles of attack, it has beenobserved that the high-energy portion of the internal flow tends to con-gregate in the upper portion of the diffuser (ref. 3) and that theexternal flow field near the afterbody is characterizedby vortex coresor lobes near the upper surface and by a thinner boundary layer on theunderside due to the effects of viscous crossflow (ref. 4). Differencesin bypass and inlet performance might be anticipated for a bypass locatedin these various flow fields. In general, however, top or bottom loca-tion of the bypass had little effect on diffuser total-pressure recovery,bypass mass-flow ratio, and hag coefficient over the range of anglesof attack and flight Mach numbers investigated in the region of stableinlet flow. At angles of attack from 30 to 9°, slightly lower dragcoefficients were obtained for the top location of the bypass. Thislower drag may be associated with the flow of the jet over the inclinedupper surface.

Of particular interest is the larger stable subcritical operatingrange obtained with the bypass located on the bottom of the diffuser fora flight Mach number of 2.0 and angles of attack of 3°~ 6°~ and 9°.This is probably associated with the effects of bypassing the internalflow. For example, the lower location of the bypass may eliminate (orreduce) separated flow over the lower surface of the internal shell,whereas bypassing air from the top may incr,easethe crossflow to thetop and thus accentuate separation on the lower surface.

Lift coefficients were slightly higher,and pitching-moment coef-ficients were more positive over the range of flight Mach nunibers,angles of attack, and engine mass-flow ratios with the bypass located onthe bottom of the diffuser, probably because of incremental lift result-ing from turning the bypass mass flow downward at the exit and becauseof an effective change in body shape due to the jet (figs. 8 and 12).

.

.

N)CDPm

NACA RM E53A29

At a flight Mach nunber of(actual angle about 0.40),obtained without bypasses,bottom of the diffuser and

E- .,..k ““”. . . .- :-*J

7

2.0 and a nominal angle of attack of 0°the lift coefficient, compared with resultswas increased 0.02 with the bypass on thedecreased 0.015 with the bypass on the top.

Other smalL performance differences between top and bottom loca-tion of the bypass exist over the range of conditions investigated;however, no other consistent trends are evident.

The schlieren photographs in figure 13 indicate that the jet fromthe bypass was discharged behind an oblique shock wave (similar to theexit flow from a sonic symmetrical nozzle), and further, that the boun-dary layer of the body had been displaced in a vertical directionbythe jet, a phenomena which was also observed in reference 5 where thejet was discharged normal to the surface. Iosses attributed to theoblique shock could be reduced by designing the bypass nozzle to re-expand to the local exit conditions. Mixing phenomena of the jet,boundary layer, and local stream are believed to be similar to thosediscussed in reference 5.

Comparison With Previous Results.

In an actual installation or application of a bypass system, the● mount of mass flow bypassed would have to vary in order to maintain

critical inlet flow over a range of engine mass-flow requirements.This couldbe accomplished by varying the minimum area of the bypassor by varying the number of open fixed-area bypasses; in either casethe sonic discharge area wouldbe a variable. Therefore, the criticalinlet flow data obtained in this investigation, with two bypasses(ref. 2), and without bypasses (ref. 1) represent three design pointswhich, considering first-order effects, define an envelope curve forthe operating characteristics of a variable mass-flow bypass system.A comparison of these data is shown in figure 9.

At the design point of the bypass (critical inlet flow,P

= 2.0,a = 00), the increase in drag attributed to bypassing 14 percen of themaximum mass flow captured by the inlet is only 20 percent of theadditive drag that would result from equivalent mass-flow spillagebehind an inlet normal shock. In reference 2, 23 percent of thecritical mass flow was bypassed and the increase in drag was also20 percent of the corresponding additive drag. At flight Mach ntiersof 1.8 and 1.6, drag coefficients at critical inlet flow are somewhathigher than those obtained with two bypasses. This apparent discrepancymay be within the accuracy of measurement of the comparatively smallforce differences. Additional contributing factors are the s~ll com-

. mutational error in mass-flow ratio and the difficulty of accuratedefinition of the point of critical inlet flow.

.

8 NACA RM E53A29

Diffuser total-pressure recoveries were about equal to thoseobtained with two bypasses (ref. 2) and slightly lower than thoseobtained without bypasses (ref. 1).

Comparisons of the thrust-minus-drag coefficients (thus includingthe net effects of pressure recovery and drag} indicate that maintainingcritical inlet conditions by means of a bypass increased the net forceon the model in the flight direction about 4 percent over that obtainedwith inlet normal-shock spillage at a flight Mach nuniberof 2.0(fig. 9(b)). Further comparison at critical inlet flow indicates amonotonic increase in thrust-minus-drag coefficient as the bypass mass-flow ratio is increased (by addition of one and then two fixed-areabypasses to the basic inlet model)”at flight Mach nunbers of 2.0, 1.8,and 1.6. This increase in thrust-minus-drag is the net result ofincreased diffuser thrust and the drag rise due to bypassing (diffwerthrust increases because the diffuser-dischargeMach numiberdecreases asthe engine mass-flow ratio is decreased). The increase in diffuserthrust is the primary cause of the increase in thrust-minus-drag sincethe change in bypass drag is comparatively small.

“

.

Application of the bypass is not necessarily restricted to main-taining critical inlet flow conditions. The amount of mass flow in “

excess of engine require~nts can be proportioned between normal-shockand bypass spillage and higher thrust-minus-drag coefficients comparedwith those attainable with nor?ml-shock spillage alone can be obtained;

,-

however, this may not be so efficient as operation at critical inletflow.

CONCLUDING REMARKS

Diffuser total-pressure recovery, bypass mass-flow ratio, and dragcoefficient were not significantly affected by vertical location (top orbottom) of the bypass over the range of an$les of attack, flight Machnunbers, and stable engine mass-flow ratios investigated. For anglesof attack from 3° to 9° at a flight Mach ntier of 2.0~ a larger stablesubcritical operating range was obtained with the bypass on the bottom.Over the range of angles of attack and flight Mach numbers investigated,the lift coefficients were higher and pitching-moment coefficients nmrepositive for the bottom bypass location. ,—

At a flight Mach nuniberof 2.0, the bypass discharged about 14 per-cent of the full-stream tube that entered the inlet with a drag increaseof only 20 percent of the additive drag that wotid result for equivalentspillage behind an inlet normal shock. Diffuser total-pressure recovery

—

was not significantly reduced compared with results obtained without abypass.

v

Lewis Flight Propulsion LaboratoryNational Advisory Committee for Aeronautics

“

Cleveland, Ohio

2E NACA RM E53A29. 9

. REFERENCES

1. Beke, Andrew, and Allen, J. L.: Force and Pressure-Recovery Charac-. teristics of a Conical-Type Nose Inlet Operating at Mach Nunbers

of 1.6 to 2.0 and at Angles of Attack to 9°.:

NACARME52130, 1952.g

2. Allen, J. L., and Beke, Andrew: Force and Pressure Recovery Charac-teristics at Supersonic Speeds of a Conical Spike Inlet withBypasses Discharging in an Axial Direction. NACA RM E52K14, 1952.

3. Esenweinj FredT., and Valerino, Alfred S.: Force and Pressure Char-acteristics for a Series of Nose Inlets at Mch Nunibersfrom 1.59to 1.99. I - Conical Spike All-External Compression Inlet withSubsonic Cowl Lip. NACARME50J26, 1951.

4. Luidens, Roger W., and Simon, Paul C.: Aerodynamic Characteristicsof NACA RM-10 Missile in 8- by 6-Foot Supersonic Wind Tunnel atMach Numbers from 1.49 to 1.98. I - Presentation and Analysis ofPressure Measurements (StabilizingFins Ren.mved). NACARME5OD1O,1950.

.

5. Nelson, William J., and Dewey, PaulE.: A Transonic Investigationof the Aero-Dynamic Characteristics of Plate- and Bell-Type Out-lets for Auxilisry Air. NACA RM L52H20, 1952.

10

Centerbody

6

TABLE I - COORDINATES

;tation,in.

-2.86-.20.1.2.3.4.5.8

1.01.52.02.53.04.05.06.07.08.09.010.012.014.016.018.020.022.424.028.032.037.146.9

Radius,in.

%al.241.321.361.391.421.451.481.561.611.731.841.922.012.142.242.312.372.422.442.462.462.442.402.322.192.031.951.751.611.W1.50

aRegion of 25‘-half-angle

NACARME53A29

Station,in.

o.015.5

1.01.52.02.53.04.05.06.07.08.08.67

cone.

cowling

Externalradius~in.

2.6712.6862.792.892.973.043.113.163.253.323.383.423*453.47

Internalradius~in ●

2.6712.6562.732.802.862.922.983.033.123*2Q3.253*303.333.35

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(o) nwlt 716W (lcdlng aft).

Figtme 3. - PhotQr@lE C& Ilypml.

(b) Jntemd view (UxMng efti).

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Section B-B Section C-C 9eotlon D-D

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Section E-B Section F-F

Fl@re 4. - DRtails Of =.ial-dlBc@% bmsa . (All ~lmens~ona1.l..ohS.)

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(a) Inletcharaoteridlos.

F&e 5. - Variatim of inlet oharacteristicm and foroe ooefflolenta with mms-flow ratio atzero ncm@d angle of attack for mnge of Mmh numbers. Mel with bream on tq.

— —

NACA RM E53AZ9

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(b) Foroe ooeffiolenta.

Figure 5. - Cmcluded . %riatlcm of inlet aharacterltiloe and foroe Oa-afllcientsWith mme-flow.

ratio at zero ncmlnal engle of att.sckfor range of Mach numbers. Mc&l with bypaan on tq.

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Mach number,

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instability

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:3 .4 .5 .6 .7 .8 .9Engine mass-flow ratio, DIJW

(a) Inlet characteristics.

FQure 6. - Variation of inlet characteristics and force coefficients tith mass-flo..fratio atnomfnal a~le nf attack of 6° for range of Mach numbers. Msdel with byp.m on top.

*U

18 NACA RM E53A29

,Mach number,

1.0 %o 2.0

8 :::---- Inlet shack

/ ~ Instability

.9 /

o“/

\u u —

o

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!&Ine mass-f~ ratio, .j~

(b) Force coefficients.

Figure 6. - Concluded. Variation of inlet characteristics and force coefficients with mass-flov ratio at ncminel angle of attack of 6° for renge of Mach numbers. Mxiel with bypsson top.

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(a) Inlet .hara.teriaticn. ‘ (b) i?oroe cc.effiolentn.

FWJ ..; j~fitim af inlet Olmmotarimtion and form Ooeffiolmtn with mm-flow ratio at nminal uqlm of attaok of 5° and 9° for Mmh. . Modal with bman m tcq .

r .“+ NACA RME53A29

wMach;umber, Angle of attack, Ref. 1

d:;

!

2.0 0

t

—.—3 —-—6 —-—

—-_1.8 ;

e1.6 ;

6‘——Inlet shook instability

.4 Solid symbols designateoritioal inlet flow

A--- = + - A b

.3

.2e

.1A

D -— - u u

04b

-.1

.6

.5 A. . -- ~ ~ ~ /,

11

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n-

c --- l-l-

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v ‘o’A ~ T ywy

0.3 .4 .5 .6 .7 .0 _ .9 1.0

Engine maea-flew ratio, m4/mo

Figure 6. - Variation of lift and pitohing-moment Ooeffloieqte with magB-flOW ratiO fOP r~a= ofMach numbers. Model with bypass on top, nominal angles .& attack, O , 3°, 6°, and 9°.

.

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.—

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NACA RM E53A29

.

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.

number,

J. ~t~-~;iSoMd8@01B deBi2Jl$3t~

aritfcal inlet flow

II ! I I I I ! I

5

.3 For ~.p.()submit Icai

‘m’”

Suparcritioalinlet flow inlet flow

~- Lnl Ill.a .2k

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Fig2

.5.4 .!3 .9 1.0

Eaglne &I-flOw r.i;io,❑4/m~8(a) Inlet characteristics.

pre 9. - Variation of inlet characteristicsand force coefficientswith mass-flowratio atm-o nominal angle of attack for ~e of Mach numbers. Model with bypass cm bottc+m.

9+

21

22 NACA RM E53A29

t

M8ch Ref. 1 Ref. 2number,%

o ●

!

2.0 —-— —--——-— —--——._ ----1.2 1

8 I& st%~lnatabi~ity--

%lid symbols designate

1.1- .I. OFltioal %nlet flow

n~ \~= I T

< .

I

$1.O0 -- ---

b0 \

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o.5 .4 .5 .6 .7 .8 . .9 1.0

Engine masm.flow ratio, m4/rno

(b) Foroe ooarrioients.

FlgLII-e 9. - concluded. variation of inlet cheraoterlntiao,and force ocdfialenta with mEsn-flow ratio at ~ro nominal angle of attack for range of Mach numbers. Model with bmaeson bottom.

NACA RM E53A29 23

.

.

co$Cu

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.

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1.0

-sFr

.6

Maoh number,

%

2.0$ 1.8

0 1.6

‘-— Ihlet shook instability

0-- -=4 . InA71A - !4

FFF73T’’’=””I A — I \ 1

-

I

.3 .4 .5Engine maes-fl; ratio, M&

.8 .9

(a) hlet charaoteristice.

Figure 10. - W.rie.tiooof Inlet cbaract%rletlcs and foroe ooeffloienta ulth uase-flov ratio atnc+nlnalangle of attaok of 80 for range of *oh numbers. Mcdel wtthbypase cm bottom.

NACA RM E53AZ9

1.0

.9

F! “8$

.4

.3

.4

.1.-.3 .4 .6 .7

&ne mase-fl.cmratio, m4&

(b) Force mefficients.

Figure 10. - Concluded. Variation of Mlet characteristics midflow ratio at ncadnal angle of attack of 60 for muse of Mach

.8 .9

force coefficients with mass-numbers. Model with bypass on

__=.

. .

—._

.

bottam.

—

.E

.

.

.

.

.

NACA RME53A29

Angle of attack,

&a;

3z 9

-- -— Inlet sbckinemilf.ty

D- --n- M l-l n

& .--- .

\

.

.

$

“~ - -.

m ~ ~

c“ u~

y.3 .4 .5 .6 .7 .8 .9

Engine Ic.9.ss-flwratio, m&,/DQ.-

(..)Inlet characteristics.

Figure U. - Verlatlon of L&t cha~cterietlcB U forck coefficlente @th mass-flow ratio atTKXAInml englee of attack or 3° and @ for Machnumber of 2.0. Ibdel tith b-s on tittom.

25

NACA RM E53A29

Angle of attack,

1.1atdeg

32 9

---- Inlet elxxkinstability

1.0 L . u

la“g L

i

iz

~- -.+ .8Q &- “

.-

0

~

~ ‘

~ “7

a;

. j

! .6 tr

LJ n

.5

A

.4

.4

a‘% n

u!?1

-..3%:

g’

o .2

I

%

.

.1.3 .4 .5 .6 .7 .8. .9

Engine mass-flow ratio, mJmo

(b) Force coefficients.

Fi~ U. - concluded. Veriation of inlet characterlstice and force coefflcients with maes-flow ratio at nminal angles of atteak of 30 and 90 for 6Kch number of 2.0. Model withb~sa on Imttan.

k~w.

.

.

.

.—

.

.

.-—.

.

N.ACARM E53A29 27

.4

.3

.2

.1

0

.6

.5

.4

.3

.2

.1

0

AN~ h

Kech Angle of Ref. 1number, attaok,

%d%

?

2.0 0 ● —-.

3 —.

6 f—-—A 9

..-—-—A-- -

---

wQ

1.8 0A

6~1.6”0

6

--- Inlet shock hetabillty

6olidsymbols designateI

critical inlet f10V

b

c –-p 4 — — - — — -u

H

1

0

.4 .5 .6 .7 .P .9 1.0

Figure 12. - Varfatinn ofof Mesh nnmber=. Model

Engine mess-flow ratio, n14jmo

lift end pltchl@-moment cmefflciante with mass-flowwith bypass cm bottm; nominal “angles of attack,0°,

ratio for a range3°, 6°, and 9°.

28 bii!i! NACA RM E53A29

(a) a=OO; ~~ .0.63; p8/F~.0.17.

(0) c4= 9°; b~ae~ bottom; n@O = 0.81; pa/P.t =0,18.

Figlz?m 1.3. : Sohlieren photograph of b~sa dim~arge at Maoh nmnber c& 2,0.

X&i

● “

NACA-Langley.3-2S.63.4oO

.

.