_. NAO_ EM A55Fi6 CONFIDENTIAL

TABLE OF COI_t_

._. Page

!. INTRODUCTION ........................ I

_ II. DEFINITIONS 3AIR-INDUCTION SYST/M ................... 3DMSION OF FORCES ............ ........ 4

_ PEEFO_tNCE PARAMETERS ................. 9: _=_ESSUHE RECOVERY 9_. . - . c , .

DRAG ............... " .......... ]_1.:_ MASS F[0W ....................... 12_ III. FEELIMIN._RY CONSIDERATIONS ................. 13

AIRCRAFT HEQUIREMENTS .................. 13J' AIEFRAME- INDUCTION- SYST_4 CGMBINATION 14-_ ENGINE- INDUCTION-SYST_ COMBINATION .......... i__. MATC_NG ........................ 16_'_ OPTIMIZATION 19_= . . . o , . . .. FLOW UNIFORMITY AND STEADINESS ....... . ..... 20' IV. DETAIL CONSIDERATIONS ................... 22_i INDUCTION ........................ 22

PRESSURE RECOVERY AND Fz_)W UNIFORMITY ......... 23Ducts ........................ 23

._ _ Area ratio .................... 23i Skin fric ion losses ................ 24

Flow sep_ation 29:% _ . , , Desi_ ....................... 27

? Subsonic Flight ................... 31Lip design ..................... 33Angle of attack .................. 36Inlet asjnmnetry .................. 37

Supersonic Flight .................. 37Supersonic compression ................ 38Limiting internal contraction ........... 40Limiting iILletMach number ............. 43Boundarj-layer shock-wave interaction ...... 45Lip design ..................... 47Mass-flow variation ................ 48Angle of attack .................. 51

DRAG .......................... 52_ Subsonic Flight ................... 53

Supersonic Flight .................. 56,, Extern_l wave drag with no spillage ........ 57_; External profile .................. 61:_ Additive drag ................... 63i Change in external wave drag ............ 64:"_, Lip bluntness ................... 65_ Net wave drag ................... 66

ii I' I _ l '_IIII llll NI , ]ll l ,l l I l , I H I _ _I I _ --'L_

1965013440-002

s/

, CO_I_DE_. :,.. ..: . : NACA_4 A55F16!i!: _ t m )1) " @ Q4 _ | e# s_ mo m_@ me

TABLE OF C0_ENTS - Concluded _!

Page

Subsonic Flight .................... 68Choked flow ..................... 68 il>_ctrumble ..................... 69T_n-duct instability ................ 70

Supersonic Flight ................... 71Causes of unsteadiness ............... 71Character of unsteadiness .............. 73Preventicn of unsteadiness ............. 74

__CE ....................... 76 i

AIRCRAFT-INDUCTION SYSTem4................ 77Effects of Inlet Location ............... 77

Subsonic flight ................... 77 {Supersonic flight .................. 79

induced Effects of _g_- of A_ _ck .......... 81Bodies ...................... 81wings....................... 8_ _-

Effects of Forebody Boundary Layer .......... 86 _Boundary-Layer Removal ................ 87 _,

Suction ........................ 88 41,

; Diversion ...................... 8_ _Submerged inlets .................. 92 k

"_ _ Combined Effects .................. 93 . iScoo_,lnc;_ementaldrag ............... 93W_Lkes........................ 94

INDU_FION-SYST_4 AIRCRAFf ................ 94 "2_,_: Drag ......................... 95

1

_ikinfriction and separation ............ 95_ausonic drag rise ................. 95

; Wave drag 96Lift and Pitching Moment 98Wing leading-edge inlets .............. 99Wing-root inlets .............. 99 Scoops ....................... I00Nacelles ...................... i01

APPENDI_ A - SYMBOLS ....................... 104_: APPENDIX B ............................ 108

REFERENCES ............................ 113BIBLIOGRAPHY ........................... 138FIGURES .............................. 171

1965013440-003

ZNACA EM A99F16 C(Sr_IDENT_ : i,8

NATIONAL ADVISORY C0MM_TTEE FOR AERONAUTICS

'; RESEARCH MEMORANDUM

AERODYNAMIC PRINCIPLES F0h _ DESIGN OF

JET-ENGIneS INDUCT!0N SYSTEMS

By Wallace F. Davis and Richard _cherrer

_" I. INTRODUCTION

/

_ An air-induction system conveys air from the atmosphere to the engine of an aircraft. Its purpose is to supply, under all flight con-

ditions, the air needed for best operation of the engine wlth the ].eastdisturbance to the external flow. In other words, to avoid penalties in

! engine size, weight, and fuel consumption: an induction system must supplyair at the maximum presto,re and with the least drag and adverse inter-

_ _ ference po'_sible. The flow to the engine must be sufficiently uniform'_ and steady to maintain engine performance _md to avoid vlb_ation and_' structural failure. The significance of tae air-induction system in

. high-speed-aircraft design has been well illustrated by SuJ_kin in refer-,; ence i. It is sho_m "_hat for fighter aircraft flying at Mach numbers le.ss_ than about i.I, the: p_-essure losses throt_h a typical normal-shock inlet!_ cause a loss in engin,_ thrust that is equivalent to less than i0 percent_:. of the wing dr_._; '_hereas, st a Mach nu_ber of 1.6, these pressure losses_ reduce the engine thrast force by an amount equal to the wing drag.

A sizable quantity of research has been dil'ected toward findingsolutions to the problems of air-inductlon system_, particularly In theMach number range from 0 to 2; but th_ results have not been consolidated

_ Into an organized group of design principles. Kuchemann and Weber have:_ written a textbook on propulsion (ref. 2) and present some discussion of_ air induction. Ho_ever, further consollda_;lon of information i_ required,:_ particularly for supersonic aircraft. It Is the purpose of this report

to assemble principles of Induction-system design for flight to a Mach;: number of e and to use existing data to show the consequences of compro-_. mlslng them. In order to accomplish thls task it _s necessary to make

an extensive search of existing literature on alr-lnductlon systems. A._ bibliography based on this search is appended to the present report. _;_i The bibliography lists "reports published since 1948 and thus extends the

blbllography of reference J. The authors acknowledge wlth gratitude the '_:

JI

/ r

j I , c _ s " _D_E_L_tAI_ o, : CA _ A55FI6

The design of an air-iaduction syst_m for an aircraft is greatly _ _.

influenced by the design of both the airframe and the englne_ and the tperformance of airframe and engine can be seriously affected by theinduction system. Therefore, the problems of air induction must be con-sidered from an ovcr=_ll vic_olnt_ az_d _. broad outli.ue must be selected _ rtc relate design principles. In this zeport, the problems of alr-inductionsystems are arranged according _o the following outline, and the principlesthat ha.'e been established for their solution are presented under theapp_opliate problem headings.

A. Definitions are presented to describe the forces involved andthe terminology used in air-induction-system design.

B. The relationships of the induction system to both airfr_ne andengina are discussed to indicate the preliminary de:_ign con-

_C_ _ siderations.

The detail design problems of ensuring high p_rformance of an '.isolated air-lnduction system and then of maintalning thisperformance when in com_blnation with ether alrcraft componentsare discussed under two beadlngs:

1. Induction, that is, the pr_ss'_re-recovery, d_ag_ flow- _uniformity, ar_d flo_-_eadiness problems encountered insupplying air to at. engine.

' ft. Interference_ or how other parts of an airframe affect theinduction system and vice versa.

This arrangement is illustrated by the following chart:

Air-induction systemsI

tvDefinitions

Preliminary co.nsideration s iI !

Aircraft requir .e_ents _tr_- , I

Airframe-i.,_duction- Engine-ind.,ction.-system combinatlun system combinstion '

I

L-tail considerations

Interference

lTessure recovery _.Airframe-induction system Ii Drag Inductlan-sy_tem airframe |i Flow steadiness and uniformity I

._, ..... ;, , ,,-....,_ .............,_ ,,,_. , ,_.

1965013440-005

!

Y_t

In order to discuss induction-systemdesign over a wide range ofoperating conditions, it is necessary to have a consistent termluoiogy.The definitions that have been selected for use in this report have allbeen used previously; and in the man_ instances where several terms havebeen used by various investigators to indicate the same concept, thechoice made here is based upon considerations of consistency, popularusage, and conven_ence.

AIR-INDUCTION SYg_EM

To define the major factors involved, consider the general arrange-ment of the following sketch:

Exlern@lsurfaces---.--- InternatIwfoces--- Streamline "_

! /_-- _o._ ._ _ _ _;_ ,o.,._ :

Lip 3

Sketch (1)

The alr-inductlon system (E_ations 1 to 3) is a part of the propulsionsystem (stations ! to 4) sac is defined to be that portion of am aircraftwhose purpose is to convey air fr_z the atmosphere to an engine. Theinduction system includes any measures taken to compress or divide theoncoming air stream that eventuall_ flows through the engine, such as theramp and boundary-layer bleed(stations I to 2) shown in the sketch.The inlet is at station 2, and the inlet area is measured in a planetangent to the most upstream point Qf the lip and normal to the mean flowdirection in this plane at maximum mass flow and zero angle of attack.If the entire cowl lip does noz lie in the inlet plane, the inlet area istaken as the area outlined by the forwardmost points on the llps projectedonto the _nlet plene. For particularly distorted inlet sha_es, these

" '" definitions are not al_ays applicable; in such cases, an area should be ,_._- . --...... f

, ....... ;......... ................ ...... -- /': .- i

1965013440-006

I/

/ i

4 : ; _ _(_?.DE_: i : ": : !NACA EM A55F16

chose_ is the most representative in terms of _nduction-systemj_erfor_ c_ Many specific definitlon_ of inlet area have been employed _ :_W-_BB-literature; two of -;_hese_Ich are particularly useful are the Iicapture area; the axial pr _jec_ionel_the _n!et ar_._and c_mpresslon- Im,_face frontal area onto the plane of station l, and the minimum cross- | T_ection area, station 2'. Each of these definitions is convenient incertaln cases,and they are identical for sharp-llp normal-shock inlets.The duct (stations 2' to 3) in the general case includes an area andshape variation along i_s length, bends_ and a plenum chamber. The _uglneintake is at station 3 and is considered to be upstream of all componentsthat are normally supplied with an engine and that are present when statictests of the engine are made. It .isthus ahead of screens and swirl vanes.The inlet lip and the fairing of external surfaces .intoother parts of theaircraft are considered to he prcblems of the induction system.

Generally speaking, there are two characteristics used to identifyalr-induction systems; namely, tDe location of the inlet on an aircraftand the method used to produce compression upstream r.fth_ inlet. Forexample, induction systems are denoted by such terms as nose, side scoop,_lug-root, conical-shock, or internal-_nntr__ctioni_l_..q:and theseexpressions _t'_combined for more complete designations.

DIVISION GF FORCES

The division of forces between a propulsive unit and nther partsof an aircraft must be carefully defined to ensure consistency. (Seeref. 4_ for example.) The air that flows through a Jet-propt_lsionsystem

i is coml_,res_ed,heated, an,_*'.benexpanded to atmospheric pressure w_+_the r_.ction from the ensuing accel_,_ationof th_ gases used to ow_.rcomethe cestr_.Inlngforces of pressure and friction and to accelerate the

i ai_c_'aft. *l_nedivision of the component forces that are included in these *i t'arustand drag forces is, to a large exteu_, arbitrary, but for practical

z_easonsspe:ific definitions must be selected. The engine designer, having! no knowledge of the airframes in which an engine might be installed,

defines engine thrust .aithquantities that are independent of installa-: tion conditions. _ne ter_ used to describe the propelling force of an

isolated engine is the "net thrust" which is the rate of change of totalmomentum (pressure plus momentum flux) of the gases handled by the enginefrom the free stream to the tail-pipe exit. The aircraft desi@uer defines jthe force available to accelerate an aircraft, that is, the net propulsive I:_ force, as the sum of _ the fo_-ces,friction aud pressure, in the flight

direction that act on all the surfaces of the aircraft (both internal andexternal) that are exposed to the flow of air. In using engine information _ .to calculate this net propulsive force, the designez must be consistentbecause it is assumed in the engine data that the pro_alsive systemreceives air with free-stream momentum, but in an aircraft installation _ . r*,this _s generally n_t so. A correction _n_t be made for the difference

_ll l 11 1 II 111 11_ : II II I I I . IIII11 r III rl 11 1 i Ill lllllll Ill " . ...... :--- I I ILl rl

1965013440-007

r_,/ ....... ,_

NACA RM A95F16 CONFIDEN_qAL 9

between the free.-streamand inlet total momentt_ in order to obtain the "_net propu_Isiveforce. The following discussion i_lustrates the considera-

tions whZeh are involved.

_:_ The net thrust force of sn engine is defined as (see Appendix A fordefinitions of symbols snd sketch (i) for the positions in_.icat_db_ thenumerical subscripts)

Fn m m4V4 - moVo + A4(P4 - Po) (1)

_!_ It is assumed in this equation that the velocity and pressure distribution_ at stations 0 axed4 are uniform and steady and that A4 is normal to the

flight direction. The net propulsive force of an aircraft is defined as

_< Fnp - (P - Po)dA - L_li - (P - Po)dA + DVex (2)in ex

Here, the pressure forces f(p - Po)dA and the viscous forces DV are ,_the components in the flight direction, and they are divided between

_i internal and exterual surfaces, Ain and Aex,, A force tending to acceler-

ate in the flight direct__onis considered positive; thus the reaction- from the accelerated gases of a jet engine causes a positive pressure

_ difference and a resultant positive force on the internal surfaces Ain._ The internal surfaces include those of the _ir-induction system (that is__ from the _tagnation point on the lea_ing edge of the ramp and from the_ stagnation point on the inlet lip to the engine intake, station 3, in

_ii sketch (1)) and the engine and nozzle passages to the exit. The external._" surfaces Aex are those in sketch (1) from the forebody nose to station

_ 1 and from the stagnation point on the lip to station 4.

_! The first bracketed term of equation (2) less the force on the ramp_'_'_. is, according to the momentum theorem, equal to the rate of momentum_,j_ change between the exit _._d the plane which includes the stagnation_ points on the inlet lip (for a three-dimensional inlet)

in

where

] 9650 ] 3440-008

/'I I _ I @@ t eO O0 @_

AI area in the plane through the entry section enc].osedby thestagnation points of the internal flow _I the llp; this planeis here assumed nn_mal to T.heflight direction, _nd flow- _|ainclination angles _e assumed to be negligibly small I

Fr sum of the pressure and friction forces in the flight directionacting on the rE_p; it is a negative force.

To u._,ili_eFn in determining Fn_, the equation for the former can berewritten as the sum of the rates 5f momentum change of the gases handled

mLby the engine between the exit and station AI and from AI to the freestream

Fn = m4V4 + A4(P4 - Po) - MI + _ - moVo (4)

From equation (3),

Fn = I_Ain(P - Po)dA - DVin1 + Fr + MI - moVO

so, substituting in equation (2)

Fnp = Fn - (.\Ey-n_oVc)- Fz -I'_ (p . Po)dA + DVe_Aex

: or _f- !

JI

According to the momen_,_ntheorem, the rate of change o_ _:.>mentumthroughl

the boundary about a d finite volume of fluld is lug, '_.._heresultantof _he pressure integr_ over the 1'tee-fluidsurface -."ithe forces actingon the fluid due to soll'__urfaces. (This statem_.nt " ,he theorem assumesste_y flow and _o shesa_ !_crce_on the free-_u_d ,"_"':--.'e.)For thestreamtube between AI and t_._f:ee Btresm,

_ PIVITM +/AI (PI" Po'dA" mV _/AT_po '_0)dA" FB " Fr :

...., .......:'?:,'",',.""':"' .:''",,.,','-,"-.... ' .- _ -- " _ ..... _l_11I

'"

1965013440-009

._:_ NACA I_4A99FI6 "CONFIUA_._ _ ;' _ . 7 "

-_

or; , AI

_ _' _J_ " mV = o (P - P)dA " FB " Fr (6)

_" where FB is the body force b_tween the nose and station i in sketch (i) :acting on the air which eventually flows through the engine, if the air-

_- induction system has a boundary-layer bleed, as in sketch (i), which pre-

._- ven_s the boundary layer frc_ the forebody from entering the inlet, MI_ would not Lnclude any of the momentum decrement of this boundary layer,_. _o FB should then represent only the pressure drag on the s_rip of.%\_ external body surface which is affected by the flow to the engine. Sub-_'i_ stitutlng equation (6) into equation (9) gives the final relat-onship

'_'/:' B!/' Fnp = Fn - (p - Po)dA + DVe x (p- po)dA - F (7)

.i_ ex o

-_,_ In subsonic flight, when the flow is neither separated nor an3_here; supersonic, the determination of net propulsive force is s_newhat s-_m-

_! plifi_d. For such conditions, the flow outside the boundary layer can be._!_. considered irrotational, and D'Alembert's theorem states tha_ for a body._._ about which the streamlines close, the component of the pressure integral :.j_.." _ in the flight direction must be zero over a bounding streamtube from the ...i. upstream station at which the flow is undist1_bed to the similar down- ._i

stream station provided, in the case of a three-dimensional body, that_ it carries no llft. l.ssuming"for ease of explanation that the external ,S flow reaches ambient _ressure at station 4 and that sketch (1) is axially"_/ sy_netric, it follo:_s that._

+ AI a

_:o,- # (p - Po)dA + (p - Po)dA --0

_: AI

_j" Restating the terms of equation (7) in smaller components

._',,,.. Fn__ Fn - (P " Po)_ " Po)a_ + _x - Po)a_ " ,;

:_" DVBI

,_,; e_ I(p-_o)dA- %,:

(the integral designated B is the pressure force on the forebody from ";_' the nose to station l) so ._

..... _._

] 9650 ] 3440-0 ] 0

8 CONFg_DENTL_JJ. : -_: o":NACA _-_!A99F16q : , _teB

,_ *e etQ 6_ 6._ oo

Fn_ = F_ - m_ex+_B (8) iwhere DVB is she friction force on the forebody 3urface that affects !

the flow to the engine. In equation (8) DVB (and in eqlmtion (7) )AI(p-po)dA-FB for the of rotational flow) is the corrective term

case

req_uired by the _finitlon of the component forces of Fnp. The enginenet thrust is the rate. _f m_entum change from the free stream to the

tm_.l-pipe exit (eq. (1)), hut part of this momentum change DVB cannotbe cha_-ged to the internal flow because it i_ accounted for in the

F.xternal flow as a part of DVe x. _ avoid the inclusion of DVB twicein Fnp, the _x.entum at the i_itlal _catlon of the internal flow must becorrected to lccal conditions, which means that DVB _ast be added intothe equation for Fnp because the true inlet momentum is less than that

as definea (_,Vo) and thus tends to increase FT.p. In the event theboundary layer from external _u_faces is removed from the engine flowby a b_Jndary-layer bleed _uch as that of sketch (1), Fu is not affectedby th_s loss in stream m_nentum, and the correction I>_B is unnecessary.Then

Fnp : Fn - DVex (9)

Taking boundary layer into an induction system does not of course, resultin only an additive correction, for Fn decreases because of the lossin pressure at the engine face _nd the decrease in m4 and V4 which _t fbe suffered by an engine _th a limiting design temperature. Eowevel-, ifDVB increases faster than Fn decreases, there can be an improvement in

Fnp as boundary layer is taken into the _nuduccion system. Quick in ',reference 9 shows that for a certain engine a decrease in specific fuelconsumption and an _crease in available thrust can be produced by takingboundary layer from a forebody into the engine at flight speeds less thanabout 300 mpn. At greater speeds, the thrust decreased r{pidly relativetc that of an engine takLng in no boundary la_er because .;fthe increasingcomp2 esscr inlet temperature and because of the loss in dynamic compres-sion 8_head of the engine. (See also ref. 2, p. 209.)

If the p__essur_ a_. station 4 is not equal to ambient pressure, then

AI 4(p /_/ (P - P)dA +/ - Po)dA + (p - Po)dA = 0o AI 4

L-_"

I

---- ..... " " "- - - 7L " "

1965013440-011

@_.;_ NACA _4 Ag.SF16 CONFIDE_IIUikL 9

_ .and

'"_ _ = Fn + _ (p - Po)dA- DVex + DVB (]0),# _ Fn-o

?

*qin other words, a correction must be made for the mc_entum chs_ige occur-

,_ ring in the jet which affects the flow and thus the forces, as p=ev-lously_: defined, which act on the system. This correction is a pressure-drag

force which acts on the external surfaces. (See ref. 6.) The fact that._: sy_ne_ry is not a necessary condition for the preceding equations for%

subsonic potential flow has been demonstrated in reference 7. It can_ also be seen fr_n the fact that if a closed body, which according _o

i_ the asst_l flow conditions can have no pressure drag, is added to theaystem, the s_etry is destroyed and the total pressure drag must still._.. be zero if the flow remains irrotationa!.r

Z_: _t_ON4ANC_ P/_tANN_RS

The basic terms used in describing the performance of air-induction_': systems are pressure recovery, drag, and r_ss flow. A description of each :_ of these concepts follows.

_ PRESSURE EECOV_NY5;

"_'_'_ Several. terms have been used to describe r,he performance of air-_: induction systems in regard to their effectiveness in providing an engine_, with hig_,-pressure air. The total-pressure ratio Pts/Pt o is the average_ total pressure at the engine intake Pt._ divided by the total pressure_ available from flight. (Methods of measurement and the determination of_ the effective Pts in non_uiform flow _e discussed in Appendix B. )_ This ratio is used when an air-induction system is being considered in

--_ relation to an engine_airframe combination because it is directly related_o the net thrust and _he fuel consumption. K_chemann and Weber show

_'_ by a simplified analysis of turbojet engines in reference 2 (p. 197) that

AFn Fni" Fna =L(I- Pt_h (ii),_ Fn-'-_= Fnl P-_o/

"_. A(Q/Fn) (Q/Fn) i (Q/Fn)a = (i - L) - (12)_' (Ql'_U)i (Q/Fn)i Pt

J__lll&

] 9650 ] 3440-0 ] 2

i0 ; _ ........... "':, : C@__AL: ..... .: : _ACA EM Ag_FI6i _ _ ii quql ii

,,. _ :l @oe s_ c,_s t9

where I

%

+ -- -'gZT-1 I"vt-qJ j

l

_j jet eificiency, Pl + (vj/vo)

Po-- pressure ratio across the engine exit nozzlePtna actual installation with induction-system losses

i ideal installation without induction-system losses

Q fuel consumptioni

_,_ _ depends on engine design and flight conditions and is greater.i than i. _ decrease in total-pressure ratio reduces the engine net thrust '

and increases the specific fuel consumption with a greater effect on thethrust reduction. This occurs because the net thrust decreases with boththe mass flow and the jet velocity while the fuel that can be burned .

; decreases only as the mass flow for a fixed turbine inlet temperature.(See _!so refs. 8 and 9-)

Ram-recovery ratio (pts-Po)/(Pto-Po) is the ratio of differences_ in total pressure as measured at the engine face _nd ambient static pres-,

sure Pts'Po and the total pressure and static pressure in the undis-turbed stream Pto'Po- This parameter is t_eful because experience ha_,demonstrated it to be only a weak function of Mach number for well-designed systems in subsonic flow at a fixed mass-flow ratio. (See iiref. 10. ) Thus, the results of low-speed wind-tunnel tests can be extra- i_polated to high subsonic Mach numbers (of the order of 0.9) for condi-_ionsin which the total-pressure profile at the inlet in flight is simulatedin the tests.l Conversion from ram-recovery ratio to total-pressureratio is accomplished by the formula:

ZSee reference ll for a discussion of equiv_lent mass-flow ratiosto be used in low-speed tests s_mulating hlgh-speed conditions. Theequivalent mass-flow ratio is one which produces the same pressure riseahead of an inlet at low speed as occurs at high speed and thus is usefulin simulating conditions for configurations which nave a boundary layer

1 growing on surfaces ahead of the inlet. _ .....

L

1965013440-013

_ NACA I_4A99F16 C(,_!D_._EI/_L _ ll

Pts " Po 7 - 1

_. pts Pt-_ Poo + _ M2 - i + i- (13)

P+. .-!_

_u (i 7.i 2)7_ I2

C_es of this variation for 7 = l.& are presented in figure i. (Through-out thls report 7 is assumed to be uqual to 1.4.)

The parameter i- [(Pts-Pto)/q2] has frequently been used to describe_ losses in duct systams. As with ram-recovery ratio, tests of subsonic_._ difusers with unseparated flow have shown little variation of this param-

eter with Mach number; but, also, it is not directly related to engine_. performance. With air-induction systems, q2 can be estimated for most

_= operating condit__ons _ithout resorting to detailed flow measurements st__ the inlet. At the _hlghmass-flow ratios which occILr in take-off, the,_' major losses in press_re occur at the inlet lips, and it is a fair asst_p--'- tion that Pt2---Pts Then, q2 can be calculated from the measured mass-_ flow, A2, and ptS . However, at mass-flow ratios of the order of i, the._ : major lo._ses occur in the d_ct and Pt2---Pto under which conditions it is'_ more reasonable to calculate q2 on the basis of Pto" If the parameter

_: is used, _he conditions for the determination of q2 must be specifically_'-. stated to a_oid confusion.

_&DRAG

!-_ The drag coefficient of an air-induction _ystem is the dimensionless "-__ ratio of force in the flight direction caused by an air-induction system2_. being added to an airframe-engine combination to the product of thea_mamlc pressure of flight and a characteristic area of the inductiont: sys._em. As indicated in the previo_,s discussion, it is necessary to be_'i consistent in defining drag; the bracketed term of equation (7), the n._t "_" drag E.,_, can be regarded a_ the drag force which is consistent with _he',_ definition of net thrust Fn usually used in computing net propulsive force Fno. _ne bracketed term of equation (7), in the general case,

'-#. includes _uch more than the drag force of the air-induction sysbem, for"f_: the drag of basic body, wing, tail, etc., must, of course, be included-_-_- in tha net propulsive force. However, for the present discuss.lon, it_-_ is assumed that only a scoop arrangement such as that of sketch (I) is_ being considered. The force on the air-induction system is the pressure __" and friction z'orc_s caused by adding the scoop to a basic body plus the"_ pressure _ntegral on the free surface of the engine-flow streamtube minu._

-,4C

1965013440-014

_ Jr

I;!

12 ":[ " CAm: t - ( I _t esl 4e t

the body forces acting on +.his stre&utube. 2 This difference of pressure 2integral and body force has been ca/led the "scoop incremental drag."(See refs. 7 and 12..) In the present devclopment, the ramp was considered ;_-_t of the air-induction system, and the force on it does not appear inthe scoop incremental drag. However, if a ramp (possibly because it is ,_a portion of a colony) is considered no5 a part of the internal system,but to contribute an external force, then the portion of it affecting _the engine flow must be included in FB of the scoop incremental c1_ag. _ :if the configuration has a nose inlet and there is no forebody acting onthe er

AS rl6 ' :., : :,, 13 "

_;_ properties are used. However, in the general case, mo is easier co '=fA DVdA, and in subsonic flow both ratios can be greater

_, : evaluate than mc clthan 1. (See p. 4 for definition of capture area Acl.)

/ 2. The mass-flow ratio m2_/m2,* is used for _he static condition_ when Vo=O. This ratio is based on the flow rate for choked flow at

station 2'. The m_ss flow, m2'*, is equal to p*V*A2, where p* and V*are the density and velocity for flow at a M_ch number o_ 1 at the pre-scribed ambient pressure and temperature. This ratio has been found tocorrelate data well, and it indicates how near the flow quantity is to

_" the maximum possible. As will be shown later, t is a criterion of the_' excellence of lip design for low-speed flight. For flight speeds other

]_.; than zero and for isentropic flow, the two definitions of mass-flow ratioJ ; are related by the equation

L_; ma' Am' ( 1 _ a_a(7.1) ._.- 0.579 m2--_. A2 + Mo

Aa, - i

"! !il which is plotted in figure 2 for _ = 1.O. The choking limit for a .

i sharp lip inlet, from reference 14, is also shown in figure 2.

llI. I._WS_L_NARY CONSIDERATIONS

_( AIRCRAFT REQUIRE_--_.,__TS

_- As discussed in reference 15, aircraft requirements are the basis_ for the choice of both airframe and engine. Since one of the considera-

tions of airframe design is that of the induction system and since the_ engine performs_ce is affected by the internal aerodynamic problems of

induction, the considerations of the air-induction system enter intothe preli_tlnary layout of aircraft; and they must be viewed from the

'_ standpoint of the flight requirements. Aircraft range and endurance,_: for instance, are dictated by fuel consumption, which is affected by the_- drag and pressure recovery of the induction system. Similarly, take-off_ distanoe, rate of cl_mb, maneuvering accelerations, etc._ depend uponi_' net propulsive force and hence on induction-system drag and pressure

reoovery. Aside from these performsnce requArements that vary w__th air-"_" craft parpoae, there are other, less tangible, requirements that must be_ taken into a_:ount in any design. For example, safety, vulnerability,

_!_ and serviceability consideration=. _ffect engine location and thus the -'_ type of air-lnduct_.on syste.mr: The emphasis on any psA_ticular requirement

1965013440-016

i14 CO_D' .a_.__. { "'" :"{ NACA EM A95FI6

fdepends upon the l;_tended mission. Thus, the /[esign of an alr-inductlonsystem must be adapted by compromises to suit many requirements invarious degrees. _

AIRFRAME-INDUCTION-SYSTEM COMBINATION

To illustrate some of the problems encountered in fitting sxl Endue-tion system to an airframe and to introduce some of the types of inlets _that have been developed for various engine locations, the progression _of design problems with increasing size of airplane is briefly discussed. _ !

C_-rrent design practice for high-speed turbojet-powered aircraft can be _indicated by the following compilation:

_ el_length Number

of x ettypeand I Xe hAirplane Engine diameter engln_=s location /_----_e diameter

F-86F 14 1 !Fuselage open nose i 9.5 1F-86D 14._ i Fuselage nose scoop 9.5F4D-I 15 1 Wing root 4.91FSU-1 16 1 Fuselage nose scoop 9FTU-1 17 2 Fuselage side scoops 6F-IO0 17 1 Fuselage opn nose 9F-8_E 17 1 _h_selage open nose 6XF-I04 18 1 Fuselage side scoops 9.71XF-105 18 1 Extended wing root 71F-89 20 2 Fuselage side scoops 2

_ F4D-2 20..9 1 Extended wing root 9F-IOI 21.9 2 Wing root 3B-97 22 2 Nacelles, open nose 1.9 ,"A3D-! 23 2 Nacelles, open nose i._F-IOPA 24 1 Fuselage side scoops I01X-3 30 2 Fuselage side scoops 3.9B-47 40 6 Nacellos, open nose 1.9B-_2 44 8 Nacelles, open nose 1.5

L

iThese airplanes have two inlets for one engine, and the ratio of ductlength to engine diameter is for a reference diameter correspondingto half the engine frontal area.

Airplane size relative to the engine is indicated by the ratio of fuselagelength to engine diameter. For small airplanes with one englne_ in whichthis ratio is less than 18, an inlet located in the fuselage nose or

! underslung Just behind the no_e has been used most frequently. From the_ induction-system standpoint, such locations are desirable because the

problems associa*ed with boundary layer fl_wlng into the inlet are eithereliminated or minimized. _e underslung inlet, in addition, maintains

_ H H ..... H

": ..... ," , , .......... :.......... l _ II _ II _ -.....

1965013440-017

_' NACA _ A55F16 C0_FIP_._ .... l_

_ performauce at off-design positive angles of attack because the flo_ isdeflected into the inlet by the nose. A_ the ratio of fuselage-to-englne

_- size increases, or if nose _-olumeis required for equipment, 3coops f_therback _n the fuselage or wlng-root inlets s_-eused. From the induction

_ standpoint, an underslung scoop position is again desirable because of_ the off-design angle-of-attack performance _d because the body bounds_y

layer is the thinnest on the windward side_ '.ThisI,o_tion has, however__ been avoided because of the possibility of foreign-object damage to engines_" during _um-up, taxilr_, or take-off,s The wing-root inlet has a possible_. advantage over _coops in that the portion of the inlet perimeter adjacent

_ to the body can be relatively short, thereby reducing the proportion of_ body boundary layer flowing into the i_let. _ku_the_._more,with m_Itiple_ engiues the ducts cau be short and the flow unimped_ by beng_so For mid-

_i wing aircraft, the wing-root inset is in a region of large induced flowangles, both from the body _ua wing at subsonic spe_s, so special pre-cautions must be taken to instu-eadequate perfor22nce at off-deslgn anglesof attack. For a high-_ing airplane, a design probl,m_of the wing-rootinlet at angle of attack i_ the thick bounda_ layer on the leeboardside

of the body. _:

_' For aircraft of greater relative size (fu_elage-length-to-englne- '__, diameter ratio _ 22) there are several possible location_ with the choice_ depending on many considerations. For engines clustered in the fuselage_

_ , scoop inlets can be used; for engines in the wlng-root or buried in the.i_. wing, wing-root, _Ing-leadlng-edge, or, for very large aircraft, under-_ slung wing scoops are possibilities. However, nacelles with a simple

nose inlet have been used most frequently. Such _a_ramEementsare deslr-__:' able from the air-induction standpoint because the ducts are short and __ straight and the problems of aircraft-induction-system interfe__nce ered_ generally reduced.

.... _.-2_GINE-INDUCTION-SYS_EMCOMBI_ATION_b

?

_ The performance of a propulsive system depends not only on the ._,_ individual characteristics of the air-induction system and of the engine,

'_! SThe studies of references 16 and 17 indicate that the flow into _.n_ airplane induction system can seldom llft damaging objects by itself. For_, instance, an inlet whose center line is two inlet diameters above _he !",_ gro_,l _d through which the flow velocity is 700 feet per second cannct_ pick up sand particles larger than about 0.02 inch in _ameter u_less a ,_

vortex forme between the inlet and the ground. However, such a vortex .'_ can form under th_ proper conditions, and if the damaging objects on the.:. ground are restrained laterally, as they would be if lodged in a crack in

a runway, the vortex will suck them into the engine; or, if objects which _: cs_ do damage (see ref. 18) are thro_u into the air by some other me_'.s, _!._' the engine can easily draw them into the inlet. Foreign.-objectdau_e to ,__', engines _s generally considered to be an operational problem, that is, one .__! of using screens, of polici_g._mPs ._ r_n:raysand of pro_r taxii_ pro- _,_ ced_es_ rather than a fact_ t location and airframe design.

9850 ] :3440-0 ] 8

wI _ I ee oe_ eo _,_o Qe

but also on the compatibility of these characteristics through the rangeof flight conditions. This problem of compatibil._tyarises because ram-

jet or turbojet engines require a specific schedule of air flow to achieve ,i.rated th_st through the flJg_htMach number and alt_tude ranges. The ._flow through a nonadjustable inlet combined with an engine varies with _fl_gbt conditions and deviates from the optLmum conditions selected for _ :the critical design point. If the range of operating conditions is suf-ficiently wide, the air-lnduction system is complicated by adjustmentsthat must be pr_ided to maintain its performance near optimum.

The general problem of combining an alr-induction system with anengire can be divided into three parts: (i) matching, (2) optimization,(3) evaluation. Matching is the determination of the mutually compatibleoperating point for an engine and air-induction system at each flightcondition; it consists simply of relating the engine flow requirementsto the air-induction-system characteristics by means of the continuityequation to determine inlet area or mass-flow ratio for prescribed operat-ing conditions. Optimization is the determination of the matching con-ditions for maximum net propulsive force or minimum specific fuel con-su_ption. This can consist of the calculation of the optimum inlet areaor mass-flow ratio for fixed system8 or of the proper variation of inletdimensions for variable systems. The two problems, matching and optimi- ._zation, are presented in same detail in the following discussion. Fv_lu-ation is the comparison of sever_! possible propuls1_.c_j_ems on an , [airframe to determine the best system for a certain mission. Evaluationscan _Ivolve many considerations in _i_tion to those of aerodynamics, _such as structure, _eigh_, mecLanical complexity, etc. However, byrestricting the propulsion-system variables to net propulsive force andfuel consumption fo_ prescribed flight plans, many valusble results canbe obtained from an evaluation study. For example, Fradenburgh andKremzier in reference 19 describe an evaluation of the effects of w_:ciouspropulsive systems on aircraft range. Another approach, which is similsr_o that used by W_worth and Kelber in comparing Jet engines (ref. 20),is to determine the allowable weights fo. the _Ast_D_lationof each of

i several air-lnduction systems on an airframe having a prescribed range.Such an evaluation provides the designer with the information necessary [to select possible mechanical arrangements_ These studies are part ofthe general problem of power-plant-aircraft optimization discussed inreference 15.

W_

MA,Ib_IKING

The problem of matching an air-induction system and an engine requiresknowledge of the performance characteristics of each, and the problem of

._ optimizing the design for a special airplsne requires knowledgc of the

i __ , _m=

i

I

i '' ' - l I --L_ .]lk_l I I ___2 " _,_v_l J .... JJ- t' ,_ 11'

1965013440-019

NACA P_ Ag_F16 COE_DENTIAL .... 1.7 7.

characteristics through _ wide r__a%e of flight conditions. 4 These char-acteristics are determined by analysis and tests, but since in the pre-

liminary stages the air-induction system has not yet been designed, its, performance _n_st be assume_ from past experience or by determining what_i performance is necessary and then striving to dezign and develop an_ _'r_em_nt that will accomplish the goal.

< To illustrate a method for matching a turbojet engine and an air-/< induction-system combinstion, the variation of corrected weight flow of: air for an engine (Wac=Wa_8/8) as a function of Mach number and the vari-..: ation of the pressure recovery of the air-induction system with mass-_ flow ratio as shown in sketch (2) are assumed to be known.

?i,

._- Wae Pt--!/ Pto

_ n,h Mo, a,_8,A2/A.'-:. Mo roW'too

_. Sketch (2)

_ For a complete analysis, this information must be available for each_' parameter indicated on the sketch; that is, the flow variation must be%.'_il known for the expected range of engine rotational speed n and of flight_. altitude h. The induction-system variation must be known for the Mach

number Mo, angle-of-attack _, and angle-of-sideslip _ ranges, and'_ posslb_y for a range of the ratio of inlet area to body fron+al area:_> _2/AM, although in the usual case changes in this ratio are small and'- their effec%s are negligible. Transposing the continuity equation

'_ (assuming uniform flow at all stations) into engine-inlet terminology by_,_ _See reference 21 for a discussion of engine performance psrameters;

reference 22 for an analysis of turbojet-engine-inlet matching; refer- ._ ences 8, 23, and 24 for relationships between engine and induction-system !_" performance and methods of determining optimum performance conditions;

and references 2_ and _ for studies of the penalties associated _Ith ""i";,: mi smatchin_. ' ........ ,'

....... Illlll II IIII I -- I .L_..]II II I I Illllll I I II .,,.]IIsL_i_II

9650i3440-020

i,f

i!

defining 8-pts/2SL, 8o-=Pto/PSL, and q-e- T_s/T-_L

mm_ peV_A._ A o _'_s

=o poVoA21_; :

gives .:{

{w_J_P_ m2A28 Pto - gPSLa_SL _ .7'+1 I ';

+ -N_ ._

(16) {= 8_.__ l+

when

7 i.4

32.] 7 ft/sec 2

PSL 0.002376slugs/ft S

ASL lll7 ft/secz

This relationship can be represented graphically so that from the kno'.mi engine and air-inducticm system characteristics the inlet area required "i

to match the engine at the _elected induction-system conditions can be; readily determined as illustl-ated in fi_e 3. Thus, for a given flight: condition o_ Mach number and altitude (sketch (2)), a mass-flow ratio

is selected and the corresponding pressure ratio determined from thealr-induction-sy_tem performance data; _he corrected engine weight flow

'. is determined from the engine curve; and the p_oper inlet area is deter-mined by the intersection of the corresponding horizontal and verticallines in the third quadrant of figure 3. This inlet area f'zrnlshes the

engine the pruper volume rate of flow at the chosen mass-flow ratio, but,this is_ o: course, not necessarily the mass-flo_ ratio that producesthe msmAmum net propulsive force or the minimum fuel conscription.

A similar method can also be used to study matching at static ccn-i ditiO_S where the _ass-flow ratio m_/mo has no significance. Defining

f

"]I _- ....,:_ ......................."_.......... _ '"'' _"'_-"...............'-

196501:3440-021

-= a I '..

_L_ J_ D |_$ m sll

NACA I]4A_GFI6 C,)N_ID_.NT!_'. : ' 19Ii

o, inlet Mach number M2' as that which would exist if the flow to station2 were isentropicj5

Wa_ Pts 894M2r- (17)

,. _ ,2_e_' A25 Pto (I + 0._a ,

; |? This equation corresponds to equation (16) if ma/mo=l and M2 is sub-_i stituted for Mo. With these changes, firo_re3 can be adapted to static "r conditions. Information on p_/pt o as a function of me/me* can be!_ converted to a function of M2' by tne_:relation

i ma = 1.728M._'(I+ O._e'2) "s (18)

me*

and this variation together wzth the }mown engine ch_cacteristlcs cad be/ used to determine the inlet area required to match the engine or thepenalties resulting from mismatching.

OPTIMIZATION> 1.8. If _he initial boundary layer is thick, the

! maximum slope cannot be large; in fact, the two slopes become equal. Theda_a indicate that a 3 final divergence angl_ on a wall, or a 6 includedangle, should be used with bo_J.hthick and thln initial boundary layers.

o These qualitative considerations indicate t_at for thick initialboundar_ layers and high inltisl Mach numbers, a diffusing straight ductshould have a faired entry section and p con _al diffuser of includedangle no greater than 8 (6 included angle plus a m_ximum of 2 forboundary-layer compensation). For other conditions, _air duct shapeswhich satisfy these considerations can be conveniently expressed as

_The data on conical diffusers from these references were analyzedto determine desirable duct shapes by selecting longitudinal pressuredistributions for which H_ 1.8, and then calc_lating new duct shapes

, from one-dimensional relationships for this pressure distribution andi value_ of M_, approaching i. The resulting'calculated shapes all hs_e

small initial slopes because, as shown by equation (23), the Mach numbergradient (i.e., the slope of the wall) must decrease to maintain a

i constant initial gradient with local Mach ru_'_r.pressure increasing

1965013440-031

NACA EM AS_F16 CONFIDFI,_IAL '- 29

_: expo_lential functions of tnc duct axial coordinate. Tests ,_ere made of a fs_uily of such diffusers with _ ratio of throat area to exit area of

1 to 2 and a variation of the ratio of duct length to t_moat diameter of_ fr_n _ to 9. Tests were made with both _=parated and attached initial

boundary layers at mass-flow ratios up to t_ maximum, and the resultsare reported in reference 64. Dat_from these and other tests are co_- 1.00

_.-Predictionsby methodof ref. 44

_ pared in sketch (6) for the condi- _ ,____:tion of an attached initial boundary _'Conical0ref. 56 I/ _'_- 8" Con_coI-O5layer. It is apparent that, for this _'-_96 L f "_._ ..,;-o t_1 olq +

_, comparison, the ratio of initial _ 13.5-O, ref. 64_'--_; boundary-layer thickness to throat 8 ---_ l 12"o Oo,'icoIaref 48_'_i irecovery than does diffuser shape. __: The measurements of reference 64 show ._ MidstreamMoth number,O.,_

- B8

i',' that the important effect of duct shape _ I I I]----,is on flow uniformity and steadiness, I:_ for the uniformity ratio VM/V varied ......:,' from 1.12 to 1.29 for ducts differing 0 .004 D08 .Or2 _6 D20

:n total-press,are ratio by only 0.02 Bou.d_y-my_r,_c_essrotM,(8/r)_oin tests with a thin initial boundary Sketch (6)

layer ((S/'r)2' =0_0014) and a hig__ initial Mach number (142' _-0.89).

Furthermore, t_ ducts having nearly 1.00 __ recoveryfor tfiinnest1,_ _qual uniformity and pressure recovery -_ unseparated, __ u.-__Y" differed by a large amount in the _ _15.5__6_quality of flow steadiness at hi@h 96,_ inlet Math numbers. The comparison

of pressure recovery predicted by 4__ 92 __. the method of reference _ with the ._, experimental measurements of sketch (6) _ .88

shows that the prediction is onlyacc_.ate when the initial boundary-layer thickness is very small. If it _ .84-----cx

' I M._%_eom Moch I

is not small, the effective skin- _ | ;_umbe,',08 flfriction coefficient is larger than _ .80 that indicated by equation (2_I) and [----_--R,eference 64|

experiments are necessary for accurate! loss predictions. (The data for

_ketch (6), and also (7), were cal- 0 D2 04 06 08 JO: c_iated according to the mass-derived Om_oceme_ th_kr_s_ ratk%(_*/r)z,

method; see Appendix B. The magnitude

of the difTerence between experiment Sketch (7)and theory depends upon which methodof data reduction is used; the

zlThe ducts of reference 64 are designated by numbers which indicetethe maximum slope in terms of incl_ed angle and the length of ez:t;-ysec-tion in terms of inlet radius. Thus_ 8 conical -0._ indicates a conical :"

divergence of 8 and an exponentz_J_ly faired entry section of 0._ inlet _'radius in length.

"_I _ ....... IIIII II L_ _ 11 IU IIII _ & . _ ,,_ I -_r 7 .... _._

1965013440-032

/3o CO_ZDENT_A,L:, : .'': ':NAOA _ A99F16ix ; 6

: te e

difference shown in sketch (6) would be smaller if the data had beenredaced by the mss-fl_w weighting method.)

Sketch (7) shows the results of tests reported in reference 64 forthree ducts w_th separated initial boundary layers. The data show thatan extended entry section increases the skin-frlctlon losses when theinitial boundary la_fer is unseparated; therefore, if separaticm in theentering flo_r can be avoided, a tong entry is umdesirable. However, withinitial separation which, as will be discussed later, c_n occur in low- _,speed flight _t high mass-flow ratios or in high-speed flight at low mass- i_flow ratios, some entry length improves duct performance because it givesthe boundary layer an opportunity to reattach. The fact tb_t _he _ressurerecovery can be higher for the long duct with the separated boun;_j layerthan with the _mseparated profile indicates that reattacP_nent occurredafter relatively extensive separation and that the small skin-f_.qctionforce in the region of separation reduc _d the over-all losses. In regardto flow uniformity, the results of reference 64 show that for short ducts _the flov is more uniform if the initial boundary layer is attached ratherthan separated. For a given initial profile of the separated type, thefinal uniformity is improved if the duct is made longer.

Reference 64 reports tests which were inteLRed to investigate to r _neextent the mantu'acturing tolerances required in duct constractlon. Meas-urements were made with a duct having d_xTerent degrees of surface rough-ness, waviness, and leakage. It was found that roughness caused byscratching the surfaces with coarse _andpaper or by putting discrete stepsin the duct walls, as could occur _Ith Joints that are not flush, had noeffect on the diffused flow. The maximum magnitude of the roughness was about 0.7 the momentum thickness of the boundary layer _t the duct throat.

The maximum wackiness tested was similar to that which would occur becausei of pressure loads in hlgh-.speed flight; circumferential stiffeners were

assumed to be 0.6re, apart, and the deflection was varied up to 19 tiuesthe momentum thickness, or 1.9 times the boundary-layer thickness, at the

! duct throat. For m_ss-flow ratios toe'/me'* below 0.89, even the maximum! waviness tested had a negligible effect on the final flow. At greater

mass-flow ratic_ 3 the maximum waviness reduced the pressure recovery,uniformity, _I steadiness only slightly. Leakage, as might occur through

i Joints in duct wslls during high mass-flow operation in run-up on take-off,was fou_ to have negligible effects when the leaks were in the low-

: velocity region of a duct However, leakage nea_. the duct inlet caused: separation with ensuing sizable pressure losses and flow nonunlformity, j

The internal-flow systems of most aircraft h_ve some offset between !_the inlet Grid the exit, transitions in cross-section _hape, and J_mctures i

i wiLh other dacts, all of which can cause losses in pressure recover[.-. .The general pzoblem in the design of these elements Is the same as that of

%1

,k,,............. _-,..,-m,m_ _' I _111 ' J

I

9650"13440-033

NACA RM A99FI6 CO_IDF_I_FI_L 3

_ a _ubsonic diffuser, tk_t is. the preven'ion of local separation and_ reduction of skin friction. I_ One design f_ture that has always been

beneficial is the use of generous fillets to a" _id angled corner_. (See_S refs. 67 and 68.) He_ever, since the factors which cause pressure losses_ differ w_th each duct configuration, it is diff__cv_ltto apply accurately_ general design inform%tion. The dat_ of references 28, 60, 61, 69. and-i_i 70 indicate the trends to be expected. The m_nitude of the total-presure_ losses in s-bends is demonstrated by the tests of reference 71. Eela-_ tively short ducts (l/rs = 4.0) with several inlet cross-section _hapes_ and a circular exit we_ze tested at a _ach number of 1.9. The inlet had

_ a wedge-shaped externsl-co_pression surface and the exit certer line of{_ the auct was offset 1.9 exi_ radii, rs, from the dnlet center line. The

_ maximum total-pressure ratios measured with the ducts were of the order_, of 6 percent less than those measured wi_h a straight duct. Reducing_ the mass-flow ratio decreased this difference to about 3 percent, a fact_ which indicatez the dependence of duct losses on inlet Mach number._] Altho_=h the total-pressure losses could be reduced by reducing mass-flow;_' ratio, the exit velocity distributions show considerable nonuniformity_ for _hese conditions. Tests with offsets of one and two inlet radii_: reported in refe-._nce 64 indicate similar results. The center lines of ;

_ these offsets were s_ooth curves similar to those of the duct-wall con-.... tours. At a mass-.flow ratio of 0.9 with a thin initial boundary layer,_ the 1-radius offset reduced the total-pressure ra_io 3 percept from that

of a straight duct, and the _-radii offset r-_aced it 6 percent. The_ steadiness and uniformity qualities of the flow deck-eased in a correspond-_-_" ing manner. For example, with the thin initial bou_dary layer, the maxi-

mum mass-flow ratio fcr steady flow was about 0.9 for the straight ductand 0.7 for the duct w_th the O-radii offset. A fourfold increase in

_ the initial boundary-layer th_._kness reduced the l'_tter mass-flow ratioto o.A. It is apparent that deviating from the optimum aerodynamic design

_. of a duct can have serious consequences.

Subsonic Flight

_, Since in subsonic flow, press'_re losses snd nommiformity result

_ from skin fr_ction, separation, and en+ering flow that is asymmetric withf respect to the inlet, the _nduction-system design problems in subsonic

%

_The design prin?_p_es for annular subsonic diffusers are like,_' those of diffusers without center bodies, but the annular type, having_ more wetted area, has l_rger frictional pressure losses. Ltudies of_ annular diffusers are reported in references 65 and 66.

"19650 ] 3440-034

t I _ *, -52 "_C"_II_lPE_f'r-AL : : ": ! :"iNACA P_4 A._._t,, , ,._ el el@ ol

flight e_e to provide conditions that avoid or minimize thesa factors.Skin friction and inte_al separation _l'e problems of duct design; the " ;problems of separation in the inlet and symmetry ea'e discussed in this ,_sect, on.

To illustrate the conditions which lead to the principP_ separationproblem of __nlet d_sign in subsonic flight, sketch ,c:_v shows a typicalcui-'e of the air requirements of a turbojet engine in terms of the free-stre-o_msrea of the engine-air streamtube Ao as a function of fligh*

12, Seeievel

SfrofosphereFlighfschedule

I0

___ 1.0

Z ; ----- .-

I

0 .4 .8"- 1.2 1.6 20Machnumber,Mo

ki

Sketch (8) Math number. It is here assumed that the airplane accelerates at sea Ilevel to a Math number of O. 8, climbs at this Mach number to sltitude, J

F

_" -"-7.... _ .....f,.':,, "" " " : ......1--_ "'_

1965013440-035

71

J

.....A _ A55F16- CO.'_IDY:I_I__AL 33

a_nd then accelerates from this cnlise condition to a Mach number of 2o_ The air requiremen_ is not only a "unction of Mo, but also of tot_!-

pres._vu_eratio and altitude, as shown, and of engine desi_--nand powersetting. Since czuising flight is usually _ important desi_, condition,

_- the irlet area A2 must be selected to produce efficient cruise perform-ence, and this, for high-speed aircraft, is g_nerally at a relatively high

_ mass-flow ratio, above sbuut 0.8. The chuiue of this mass-fLo_" ratio isa comnromi_e he_ween reaD_irements for other flight conditions p_ud thecor.riicging interests of the intern_l and exterr,al flows. A low mass-flowra_io (me/mo=Ao/Ae

e eo_

v,_ a D reat e_ @ o4_ 40 OOe @m

:'" N/1-// 'i, !//; i

1___ 1 I __1720 l .2 3 .4 5 6 _? 8 9 LO

Sketch (9)

i

in sketch (9) (Losses in the duct behind the inlet can be added to these, total-pressure ratios to determine the pres=ure at an engine face pt3.' At high mass-flow _'_tios when the lip is stalled the duct losses are :

small relative to those due to flow separation at the lip and are seld_known. ) If the inlet area is ro_ ",ctedfor the altitude, cx-aise conditionand information similar to that o_ sketch (9) shows that the mass-flow

ratio me/m2* is about 0._ ._n _ake-off_ the total-pressure ratio Pt2/Pt ,-) at the inlet is then less than 0.9. Such pressure losses correspond to ai i_- to 20-perc6nt loss in englue thrust which, of course, represents

serious limitation on _he acceleration characteristics of an airplane.The flow nonuniformlty which accompanies the total-pressure losses caneven further limit engine operation. If a smaller inlet area were chosento suit more closely the requirements of supersonic or low-altitude high-

' speed flight, the losses would be even greater. On the other." hand, the, effects of increasing fli_t speed are rapidly alleviating, i

These large pressure losses at low speeds that result from a sharpJ-_ lip can be avoided by several methods. A curved internal lip _)rofile

which the flow cam follo,_ prevents separation and the attendant nonuni- '_formlty at high mass-flow r_tios, or, for a given lip profile, the losses can be reduced by decreasing the mass-flow ratio either by increas'ng theinlet area or by taking air in through another inlet. Tests of llp _.-

' profiles on circular nose inlets at low speeds are reported in refer- _

ences 72 to 79. Some of _,he rest_Its, in terms of pts/Pto, are presented

I

18 5013440-037

NAOA RM A_FI6 CO_'D__._.,.A.'_ _5

in figure _ and are compared with the prediction of Ptm/Pto for thethin lip of sketch (9). Duct losses have not been subtrp.cted from thetheoretical prediction because a wide variety of duct designs are ccmpsmed,and, in most cases, duct losses by themselves wero not measured. For thecases in _hich smooth, nearly straight ducts were tested, the agreement

between Pts/Pt and Pt2/Pt is good at zero forward speed. However, fthe los_es for the conical-shock _uleL from reference !_ are considerably

in this particular test. The scatter of data at the maximum mass-flowratio is considerable, and a large part c_ _t is undoubtedly due toinaccuracic_ in total-pressure measurement. Blackaby _ud Watson _ref. 72)point out that near choking the flow through ducts is very unstesdy 2 and_as mentioned in Appendix B, measurements of pressure recovery by normalmethods under these conditions are not _eliable. The data on the F-8_Fand F-lO0 airplanes are from full-scale tests. The fact that they cor-relate with the data from model tests "ndicate that the effects of scaleare small. Also, since the predictions of the momentum analysis whichhave no relation to scale agree so well with experiment, negligible scaleeffects in regard to lid losses are to be expected.

The tests of reference 73 indicate that for a reasonable variationof shape external lip profile has practically no effect on internal flowoAt zero flight speed, the data of reference 72 show that pressure recoveryis not highly sensitive to internal profile, for there was little differencebetween ellipticel and circular shapes. However, as shown in figure ,internal lip profile is impnrtant at higher flight speeds, for the ellip-tical shapes are better than the circular ones. At the flight Mach numberof this figure, 0.33, a sharp lip causes relatively large losses at highmass-flow ratios, as at zero forward speed; but, in this case, the pre-diction of pt2/Pto is greater than the measurement of pts/Pto by 1 to2 percent, whereas at zero forward speed there was no difference betweentheory and experiment for high mass-flow ratios. The desirability of theelliptical profile is furthe__ substantiated by the recommendations ofPendley, Milillo, and Fleming (re._. 76). An elliptical internal shapewas selected for this investigation from previous experience, and it wasfound that the profile resulted in high total-pressure ratios for a noseinlet at zero angle of attack in the Mach nt_zber rs_uge from 0.6 to 1.1.At these flight speeds, the mass-flow ratio of an induction-system-enginecombination rapidly decreases to values less than 1 (see sketch (8)), sndthe problem of internal separation from the lip disappears. In fact, evenfor a perfectly sharp lip, sketch (9) shows that internal pressure losses ..resulting from li_ separation at the mass-flow ratios of interest (up to0.9) are small at flight Mach numbers above about 0._. Thus, at highsubsonic speeds, skin friction is the major source of pressure loss inwell-designed systems.

Some tests have been made of qrhemes for reduci,_ the mass-flow' ratio in low-speed flight to avoid lip separaticn. These methods consist

_1_. ;,_,

i

1965013440-038

i36 COffF]_NTIA_',. i : -: : _ACA f_MA55FI6e ,on 6_ 0.0 ot

oi" increasing the area t_o_h _'hich air can flow into the inductionz_v'.bern. In reference 77 a sharp-lip nose inlet was tested with a secondaryteoop having sharp llps that opened into the underside of the duct a shortdistance behind the inlet. At zero flight speed, it was found that the

variation of Pts/Pt with mt/mt* (where mt is the mass-flow throughthe total area) was nearly identical no matter how much area (up to 68 per-cent of +_o+ _? +_ _o_n ....._-_ _" j_ ".'as__,,_.,.._,_--^---"_-_ in the _uxiiiary _ccop. --Tnus,the _rovement in pressure recovery that can be expected with this method Eis entirely the result of reducing the mass-flo_" ratio for a given engineoperating condition. In reference 78 a supersonic conical-shock inletwith a sharp lip was tested with a translating cowl; that is, a shortlength of cowl including the sharp leadir_ edge could be moved forwardexposing a gap with a rounded lip and increasing the minimum throat area.Since the curve of total pressure ratio as a function of mass-flow ratiorotrot* (mt is here based on the increased throat area) for the extendedcowl lles above that with the cowl retracted, it is evident that thismethod not only increases the available inlet area, but it also improvesthe quality of the flow.

Angle of attack.- The flow approaching _a inlet can be asy_netriewith respect to the induction system axis because of the changing attitudeof aircraft for various flight conditions, because of the induced flowfield of the aircraft, or because the inlet is distorted by configurationrequi.'ements. The ultimate resttlt of such asymmetry is internal _eparatlon.Data from tests of circular nose inlets at angle of attack and a flightMath number of 0.24 (ref. 79) show that an inlet with blunt lips maintainshigh total-pressure ratios and uniform flow to greater angles of attackthan one with sharp lips. Fox" example, at an angle of _ttack of 19 and

i a mass-flo'_ ratio of 2.0, the inlet with an elliptical blunt lip attaineda tot-__l-prLz_ure ratio of 0.97 whe_-eas one with a sharp lit attained or_ly0.90. The corresponding deterioration _n flow unifoz_!ty was a differencebetween maximum and _tuimum total-pressure ratios in the duet of 0.08

; for the elliptical lip and 0.16 for the sharp lip.

At Mach numbers from 0.4 to 1.1, the .esults of references 23, 76,and 80 show that even _th shal-p lips pressure recovery is nearly insen-

, sitive to attitude to angle of attack of about 8 to mass-flow ratios ashigh as 0.9. At higher mass-flow ratios this range of insensitivity ._decreases. The sharp-l._p inlet of reference _3 suffered greater losses

: at high angles and mass-flow ratios than did t_e blunter lips of the tests;at a Mach number of 0.9, an angle of attack of 12, and a mass-flow ratio

' of 0.9 the total-pressure ratio was 0.92 whereas a blunter, but stillrelatively thin lip, had a total-pressure ratio of 0.94. For these flight

_ conditions, the mass-_lo_-ratio (m_/mo) at _ich choking occurred _,_th

the sharp lip was 0.9 and that of the blunt lip was 0.9_.

The sensitivity of an air-induction system to _g] _f sttauk _: not' only a function of llp _roflle, but it _s also affecued , +L_ _ivergence

. ':_ " ' ' . . i ". : ..'

....... i_o - : ....... _._---_"_,,,_:e_"'_ ' ,,, _"| |_i'

1965013440-039

%w

_ NACA RM A_gF16 C0k?UDE_FIAL ,". 37

of the flow behind the inlet In the _ests of reference 76 it was found_ ' that an NACA 1-40-200 cowl was more sen._itiveto angle of attack and mass-

flow ratio than a longer cowl, NACA 1J'O-;_O0,because the duct in the% shorter cowl expanded more rapidly, i_us, some lip bluntness and slow_ divergence of the flow behind the inlet p_ovldes high pressure recovery_ over a mifficient angle-of-attack range for most purposes. For a stir

greater range of insensitivity, the lower lip can be drooped and staggeredas suggested in reference 76 and tested in reference 81. In the latter

_i- investigation, a blunt, staggered-lip inlet was tested eta Mach numberf of 0.14, and it _intained high pressure recovery throughout the range

of the tests from inlet velocity ratios of 0.6 to 2.2 and angles of attack/ from -5 to 12.L

5_ Inlet asy_netry.- An inlet that is distorted relative to the axis of_ an air-induction system can .hacelarger pressure losses and greater flow_ non,Jniformitythan an axially sy_netric inlet. For instance, Seddon and

Trebble in reference 82 report tests of a wing-root inlet at zero forward_. speed. In comparing an _let swept back _2 with an unswept inlet, it

_/ was found that the losses and flow nonuniformity were about twice those t_- of the unswept inlet. A"_eadditional losses were due to separation in _#,_ the outbo_axlcorner of the inlet which resulted from the fact that, for s_, this operating condition, the flow must turn through a large angle to S_ enter the duct, since it approaches nearly normal to the inlet plane.

Guide vanes ali__._with the duct axis in the outboard portion reduced the ...._ flow nonuniformityj but increased the pressure losses. Slots in the inlet_ lips similar to wing-leading-edge slots, but not swept, reduce_ both the _-_ losses and nonun_formity because they increased the inlet area and bled_ high-energy air _zntothe region of potential separation._v

e An ._mportanteffect of inlet frontal shape is ahown by comparison of J_: the flow-6,istributionmeasurements of references 83, 84, and 8_ from tests_ of _ing-root inlets at Mach numbers from 0.6 to 1.4. The results show

_ that the uniformity of the flow in the portion of the inlet which was _'_- unaffected by the fuselage boundary layer - the outboard portion - "_as -__,_ greatly improved as the shape was changed from the acute 8_le of a tri- __ angular inlet to a semielliptical or semicircular inlet. _

_ Supersonic Flight ::_. !_

-_ The considerations of pressure recovery in sopersonic flight are !'__ more complex tk%u those at subsonic speeds because in supersonic com-_% pression of engine air the pressure losses and flow nommiformlty can be _

caused by two additional factors, shock waves and shock-wave-boundary-layer_ interaction. These factors become increasingly important as the local _"_ Mach number at which they occur increase_ above 1. Moreover, the necessary __ increase in thrust of air-consuming jet engines with speed depends upon _%_ t_e increase in tc _al pressure

//

38 @0_LDEN_AL : ..A_AEM c _ ". t _. c : : -I_v'_

m_

Pto = Po (I + 0"2No 2)s'5

ann density

Pto = po(! 0._o2)2"5Little of the available pressure and mass flow can be lost if an engine isto overcome the large dr_ forces of supersonic flight. In many cases, themargin of excess thrust at supersonic speeds is reia_ively small, and thethrust-available and thrusl-required curves are slowly convergent. Then,small losses in total pressure cause large reductions :n acceler_.tion andmaximum-speed performance. i

I

Supersonic c_pressionZS. - Since the local Mach number at the intake !of present-day engine_ must be subsonic, the flow to the ermine of a super- I_sonic aircraft must be decelerated through a Mach number of i. Ideally, Ithis compressicn of the air can be accomplished isentropically through c lreversed Laval nozzle with no external wave drag as indicated in sketch _(i0) ; practically, shock-free internal flow canno_ be attained because i

/ I

intet'nolcompresslo,,througaM_@waves Extecr_coml_ssio,t_mug_shock,ayes

Normal_ compmuio, ,ntmnr:co_sskm ttwou_shockwaves

[externalendinternalcompression

(Io) - -SketchISFerri in reference 86 and Lukasie_Icz in references 53 .and 87 dis- .

_ cuss many of the principles involved in supersonic compression. In this -report, these principles are mentioned only briefly, and the emphasis is

on presentlng information that is useful in design and in point.tug out].imitations for the flight conditions under consideration-.

'/ e,_

1965013440-041

NACA A F!6 39

the flow through such a channel is in a state of neutral equillbri_n. Anydisturbance which causes a loss in total pressure between the entranceand the thrcat causes _ deczease in mass flov through the throat becausehere the area and ve _city are fixed. Air _mst then accumulate because

_ more flows into the p_sage than can flow out, and a normal sh_.k w_ve .::- is fozunedwhich must move upstream, continually growing stronger3 until

it is expelled from the channel and spills the excess aic. The shockwave caz,not _.'e-entertilechannel unless the throat is opened sufficiently

r

to pass Ch_ full mass flow at the stagnation resstu_eexisting behind the 5!normal shock wave in the free stream. (For detailed discussions of thesephenomena see refs. 86 through C9.)

'L It is, of course, not necessary to attempt supersonic compression_ either in a closed charmel or isentropically. The flow can be deceleratedv externally s_idthrough discrete shock waves as shown for several possible-.

arraz_Eementsin sketch (lO). The crudest method which entails the greatestlosses is to _ccept a normal shock wave at the free-stream Mach number. -i

__ Sine_= these normal shock losses can be redaced by decreasing the Mach. number at which they occur, hlgher tote1-pressure ratios can be attained ::-" by placing an inlet in a region of substream velocity on an aircraft, a_" will be discus.Bedsubsequently under !NTERFE._CE, or by creating obl_que_' shock waves to reduce the local Mach number but with less loss than that

of a single normal shock wave. For a given local Mach number ahead of an/ air-inductlon system, the question arises as how best to uti__ize oblique_ shock waves. Oswatitsch (ref. 90) has shown that the maximum tot_-- ..... :" T

pressur_ ratio of a two-dimen_ional I0_' multishock system occurs when the# total-pr6ssure ratio across each

oblique shock wave is the same For_o " 3

such conditions, the variation of ; _ _ k_

_ . total-pressure ratio with Mach number _-_I__ ,_ for shock-wave compression (n oblique .o_" S

waves plus termina& normal shock wave) _ .6 k _is shown in sketch (11). It is appar- _ _2< ent that the losses _rough a _ingle _ 4 .... _

normal shock wave rapidly become _._ Normal_' _. \% intolerable above a Mach number of \._ about 1.6 and that large improvements _ sho_w_, can be made by utilizing oblique 2 ,,_

i _ hck waves'_" _O

The variation of total-pressure O0 I_ Z6 _4 42ratio with deflection angle for various _. _'

approach Mach numbers in two-dimensional '._. ,

, flow is shown in figure 6 for a two-shock Sketch (ii) _system (one oblique a_.da terminal normal shock wave) and in figure 7 for ....._.a th_-ee-shocksystem. Figure 8 presents these variations for a two-_hock _','J)!

i4Detailed information and design charts on shock wa_es can be __;_obtained from such references as 91 and 92. ....._._""_

mR"_ .'_. i,_;_'

--- II llllllI -- - ._ l l__ I._ I ._ Jl.JL . . .- I . I J_lIIII I ';' _ -"

_"_ _",W_ ....... . ..... _, _ _"' ' '-i _ -- _-_-.-_-1 .... _-_--

1965013440-042

!4o :2:i. : NACA RM A55F!6 _system in conical flow _d is ts_ken from reference 53 where it is assumedthat the normal shock wave occurs at the average of the Mach number behindthe conical shock wave and on the cone surface, (Ms+Mc)/2. This assumptionis adequate for the Math nt_nher and cone-angle range of interest in theflight conditions being considered in this report because the differencebetween Ms and Mc is sls_ll, less than 0.01. It is apparent from thisfact that the maxi_dm total-pressure ratio attainable in two-dimensionaland conical flows is about the same. Lukasie_cz in _eference 53 showsthat 5his difference in total-pressure ratio at Mach numbers less then 2.0is less than 0.015. The curves of figures 6, 7, 8_Id 8 show that total-pres**_ureratios near the maximum can be maintained for a relatively widerang_ ef flow deflection angles, an important fact because an eagle can beselected which produces nearly maximum recovery at the high-speed conditionwith little decrease from the maximum possible for a considerable range oflower Mach numbers. Also, the angle can be chosen so that a detached shockwave occurs only at a low supersonic speed where the entropy rise through anormal shock wave is small. For example, at an upstream Math number of1.8, the maximum total-pressure ratio with a two-shock system is 0.945,and the corresponding flow deflection angle is 14, for which the detach-me/It Mach number is 1.57. If a lOO deflection angle were selected, onlyO.O1 would be lost in total-pressure ratio at the design Mach number, butthe shock-detachment Mach number would be reduced from 1.57 to 1.37 and,in this Mach number range, recovery would be improved several percent.The total-pressure ratios decrease beyond the maximums (the values plottedin sketch (ll) for the two-dimensional cases) because the Zosses throughthe oblique waves exceed those through the normal wave until finally theoblique wave detaches from the deflecting surface and only the pressurerecovery through a single normal shock wave is possible. The high levelof total-pressure recovery that can be attaine_ by conical-shock compres-sion has been verified at Mach numbers to 2.1 in references 13, 93, and 94.In reference 94 a center body contoured for isentropic compression at aMach number of 1.85 produced a total-pressure ratio of 0.967; with threeoblique shock waves, the total-pressure ratio was 0.954; and with two, ._

it was 0.945. In all cases, a uniform flow was measured after diffusior.These values are very close to those obtained by adding the predicted shocklosses to the experimental duct losses described prevlous]y.

Limiting internal contraction.- For internal-compression systems. through shock waves, the problem of flow stability exists as in the

reversed Laval nozzle because of the two possible stable positions of thenormal shock wave, ahead of the inlet or downstream of the t.%roat. However,at the expense of complication, this disadvantage can be overcome, and thisform of supersonic compression hes the advantage over external compression

' of deflecting the flow toward the system axis rather than away from it. Thefrontal area, external drag, and amount of turnLug in the duct can thereby .be reduced. Thus, the optimum sa.rangement for any specific case requiresdetailed evaluation. The relation between contraction ratio, total-pressureratio, and Mach number is

' C--

t

1965013440-043

i NACA RM A_F16 CO__%L' 41 _

"_ A2' Me 4 Ma "2 _ p _ |]. + Ma2 Pt2 ' .9

This relation is plotted in sketch (l,2a) .e \_ _ ._._, for isentropic flow to a Mach number of_; 1 at the throat. Also shown is the con- _I_ .7 ..... _ ..:_ traction ratio which permits isentropic & \ _,,,f.ee-_i flow to a throat Mach number of I fr_n "g

the total pressure existing behind a _ '_ \"_. normal shock wave. This is the con- " \'_ traction ratio at which supersonic flow _ _-_,Nt- .. _ $_ can be established in a fixed internal- \

_j: contraction inlet at a given flight Mach \number and is designated _start" Total- .4_" pressure-ratio curves for two positions _

j_ of the normal shock wave for _start are _ :_. also shown for the cases where the normal s .._ shock wave is at the throat and in the free _ .,

stream. It is, of course, possible for the .z,_ normal shock wave to be downstream of the _ _4 LS z.z z6 _o._. throat, in which case the pressure recovery:_: decreases _oward the lower curve in Skc.tch(12a)._. sketch (12b). It is apparent that the

_._ starting contraction ratio for a Mach ,.o_, (_,_/

_ number of 2.0, for inst_mce, is less than ;__'_ j \/_/_,%_

_ that permissible at a low(r Math number. 9 I

_ Thus, if an aircraft is to reach a Mash ._f.es(___number of 2.0 and maintain the total-

_ pressure ratios (Pte,/PtO)_start or higher_ .e_-- _ '%. the contraction ratio must decrease with . k_i. increasing flight speed above a Mach num- =_"_ .... \\_._ ber of i. Also, it is apparent that above_ a Mach number of about 1.8, the total-.L_L pressure losses with _start are unac- .e---- _.: ceptably large, and it Js desirable to _ X_':" decrease contraction ratio and incr_:ase _ .5_=--_i_, supersonic _ompression toward the isen-_ tropic value. If the throat area is ____/

adjustable, this can be done as long as .4_ _-_i the flow at the throat is supersonic._i_ For a given contraction ratio the Mach

number at the the'oatcan be calculated .s_ from equation (2_), sad the maximum

total-_ressure ratio possibl.eis _hat"21.0 1,4 I_ 2.2 Ze _ "_

_' of a normal shock wave occurring a_ -_(', M. :_Mach number .Me_ with Pt2 _/Pt2=I"

_ ...., Sketch (12%) _

m =m J_ I_1 I ..................... II II I I1111111I_111 I]1 ........... .... '/

' | ""' I

1965013440-044

t/

i eta e-42 coNfiDENCe.'. .": .: : :NACA _94 A95FI6..... : :.-[ t _ ce eo eo

However, if the flow at the throat is subsonic due either to a contractionratio that is too small or to the inlet being too larg_' for the engine-air requirement, a normal shock wave _Jaead of the inlet reduces the total-pressure ratio to t_at of the lowest curve showD in sketch (12b). Infact, this type of air-induction s_stem is sensitive to flow changes, andclose control of bOth inlet-area and contraction ratio are necessary if itis to operate with an engine through a wide range of flight conditions.The pressure recovery car. decrcasc abruptly from the _ximum possible withsmall changes in either mass flow or angle of attack (see ref. 93).

An induction system in which both inlet and throat areas were adjust-able to match engire-air requirements and provide maxim_n total-pressureratio with -"_*_ 91 contraction through two oblique shock wave,_ and aterminal n ,ave has been reported by Scherrer and Gowen in refer-ence 68. I ,_b found, as shown by the data poJmts in sketch (12), thatin this particular test a contraction ratio well below _start could bereached, but there were no significant improvements in corresponding total-pressure _-atios. It was concluded that the increasing supersonic compres-sion was counteracted by increasing losses in the duct and that greaterrefgmement in duct design was required.

Other methods than adjustable passage walls have been invest_ ,ed for ---. avoiding the flow-stability proble_ of internal-contraction inlets. Evvard and Blak_y (ref. 9_) tested an open-nose inlet in which the contracting

passage was perforated _o penmit the escape of excess flow between the

_J

< _CA _ A59FI,5 C0_DE%_ZAL 43

}_ circumvented by use of a leadi_,-edge flap on the compreasion surface. :_1 (See ref. i01.) Deflectlon of this flap toward the body reduced the_. pressure rise across "_he oblique shock wave at a given Mach number, and

delayed boundary-layer separation to lower Mach ntunbers.

i_ For the conical-shock inlet, internsl contraction can be used to'_. produce additional supersonic compression, but at the expense of encounter-_ ing the flow-stability problem and additional duct losses L_kaslewicz_ derives in reference 53 th contraction ratio _start that can be usedb

with conical-shock inlets, based upon the assumption that the entrance._i: M_ch number is the average of that behind the shock wave and on the cone !_'_ mu_face. This variation is presenc_,d in sketch (13). It is seen that for :

:._ large cone angles the permissible contraction is small. Experiments at._ Mo =1.85 (ref. 93) show that for an inlet _th a straight lip (not cam-

bered to meet the local flow), internal contraction reduces the optimumcone ang)e for maximum pressure recovery IDO

_" to about 25 as compared to 30 for _n __k_] inlet with only conical-shock compres-

,{ sion, (fig. 8). However2 the differene e 96 __._ in maximum _ossib]e recovery is small. \:_ Only for SL_Lll cone angles where the .92

_, oblique shock wave is not being fully 30"_: utilized can internal contraction 88 --

;_ produce any great advantage. Tests _ \_ r_ have been made at a Mach number of 1.85 *'.84

'! with conical-shock inlets having internal _i____

_ contraction and a perforated lip to pro- 20"_i vide :"low stability. (See ref. 94. ) .80

The results indicate very high maximum; total-pressure ratio, 0.95, for this .76,,_ arraugement. Both drag and pressure-

_ recovery measurements were made for a .7210 _____' conical-shock inlet with a 20 cone . 1.4 18 2._ 2.6 5.O_', and a perforated cowl at Mach numbers Mo

_, of 1.59, 1.79, and 1.99 in reference 96. Sketch (13)_ The resulgs indicated that even thougJa_ high pressure recovery was obtained at zero _ngle o_" attack a r_:latively_ large increase in external drag occurred rela'i;iveto simil_r unperfor_ted /.!_ inlets. The pressure recovery was relatively insensitive to mass-flo__. change above the mass-flow ratio at which shock oscillation occurred, o-__: With increasing angle of attack both the range of mass flows for steaay_' operation and the pressure recovery decreased at all M_ch numbers, the._ latter being a more pronounced decrease than with simil_r unperforated

_,' _nlets. :'i!

,_ Limiting inlet Mach number.- Fc,r extern_l-cmmpresslon systems there -_;_is no problem of flow stability as, there is with internal-com_ressi_ ,:_systems. There is, however, a li_Itation on how nesmly isentropic the _

i compression can be, or, in other _ds, on the number of oblique shock "i_

g-

196501:3440-046

44 ' _iD_.WA._'. " .'. _ : NACA RM A_-F3.6' ; .: .....

waves which it is practical to use. This limitation 8a.ises because thelarger the number of shock waves, the higher the subsonic inlet Machnumber and the greater the duct losses. Hence, _otlmum supersonic ccmpres-sic, requires excellence in duct design. The _o. lowing +_ble sho_'s thelocal Mach number and total-pressure ratio after the terminal normalshock wave in a pattern arranged wlth n oblique shock waves to producethe maximum supersonic compression at approach Math numbers of 1.5 and 2.0.Subta-acted fk-om these total-pressure ratios are the duct losses corresponm-i_ to the inlet Mach number ss measured with a duct w_th very small lossesin reference 64. Thus, for these conditions, which are probably about the

t "Duct13._ --o, (o/r_2= o.oo143Mo = 1.5 Mo = 2.0

!_t_p;_-_t__t._ _t_pt_-_t_p_

n"_-_o Pto_o _ _ _t-'_0 0.70 0.93 0.02 0.9l 0.98 0.72 O.O1 0.711 _ 98i o_ _ 74 _ o23 -94 1.00 .0_ .96 .90 -97 .03 .94

best that can be expected in the present state of practical design knowl-edge, little ca,. be g_tned by using more t'han one obllqae shock wave _.ta Mach number of 1.5 or two oblique wave_ at a Mach nz_ber of 2.6. Ifa poorer duct is used, say the duct "_Ith a thick initial boundary layerand a two-radli offset as described in reference 64, the following resultsP_e obtained when it is csmblned with shock-cumpression inlets:

Duct13.9o - _ [Offset= 2r2),(e/r)2= 0.0156-_ = 1.5 -- }YO 2.0

_t__t__-_t_pt_ _t_pt_-pt_n Me -- - M_ Pt PtoPto Pto !:Pto Pto

&

0 _,17oi0.931 o.09 0.84 0..58 0.72 0.06 0.661 ._i .98 .14 .84 .7% .90 .i0 .802 ._i .99! .16 .33 .83 .99 .i3 .82

_94 .17_.00 .83 .90 .97 .l_ .82

-_re, the advantages of high supersonic compression are further reduced.a )_ch number of 1.9, a normal shock wave might as well be used, and

_7, a Mach number of 2.0, a slx_le oblique shock wave very nearly producesmaximum pressure r-_covel_. 0swatltseh esteblishes this point in refer-ence 90 by considering the arrangement of oblique shock waves which wouldproduce the maximum static pressure behind the terminal normal shock wave.This would be the best initial condition for a poor duet installation.

"I l I

1965013440-047

t -

-f

i c is shown that oblique shock waves pz.oduc_ no improvement to a Machntunber of 1.6 and that a sngle oblique wa_e is sufficient to s Much numberof 2.0.

.At flight Mach numbers greater than 2.0, another limit appears onthe n_nber of oblique shock waves that can be us_=

": "{:CA46 .... ; C0_F_$:_. ': :-': : :- @ .t , : ' :.:

amounts of separation with subsequent reattachment are not necessarilyserious, and information is required o_ the allowable tolerances forregions of separated flo_:.

With air-inductloa systems, the shock waves that interact _th aboundary layer can originate from a change in surface slope, from neigh-bcring surfaces, or frc_ the normal shock wave which terminates supersono.compression. Bogdonoi'f._udKepler (ref. 10_) indicate that for local Machnumbers through 2.0_ a static-pressure-rise ratio of about 2 causes separa-tion. Gadd, Nolder, and Regan (ref. 106) show a value of 1.7; Nussdorfer(ref. 104) suggests a value of 1.89; Lu_asiewicz (ref. 52), Seddon(ref. 103)3 and Dailey (ref. 108) suggest 1.8, the pressure ratio acrossa normal shock wave occ:_ring at a M_ch number of 1.3; and the c.'iterionof Nitzberg aud Crandall [ . . z" (Usep/u_mltial) - 1/2] corresponds to a static-pressure-rise rctio of 1.7 (ref. 109). Such differences are due to themethod used to determine separation and to test conditions. Nussdorfer'scriterion of statlc-pressure-rise ratio of 1.9 was derived from a study ofair-induction-system data which included both :?iaQeand con:ca/ compressionsurfaces. If this criterlou is used as the one appropriate +o presentdesign methods for the case where a uo._na:shock wa:e interacts _th aturbulent boundary layer, the limitations on shock compression becauseof separation are those superimposed on the curves of total-pressure ratioas a function of flow deflection angle and Mach nl_mberpresented in fig-ures 6, 7, and 8. If it is assumed that the degree of separation at theboundary determined by Nussdorfer's criterion is sufficient to reducein_uction-s)stam performance, it is evident that in the Mach number range

up to 2.0 inlets must be designed for nearly the optimum shock confi_Ara-_ tlon. If a smaller de_lectlon angle is used, the termin,_lnormsl _hock

wave is intense snoagh to cause separation. This interaction undoubtedlydecreases oerformance in cases where the boundary layer just ahead of the

normal shock wave is on the verge of sep_ation and where the subsequent: flow is not given an opportunity to reattach. For instance, the skatch

_u f!gur_ 7 shows a condition where the pressure rise in the vicinity ofthe oblique-shock reflection could be sufficient to cause local separation

!or at least disturb the boundary layer sufficiently so that the terminalnormal shock wave would erasureseparstion. The limitations for avoiding

seperstlon in this case are more severe than indicated in this figure.Comparison of figures 6, 7, and 8 shows that a strict requirement ofavoiding bow-shock wave detachment and separa%ion due to the termiual

_ normal shock wave thro_gh a range of flight Mach numbers makes systemsin wlqlchthe configuration can be varied necessary at Mach numbers aboveabout 1.6 in two-dimensional flow and above abou_ 2.0 in conical flow.(Other reasons for variable systems and information on those that have .been tested will be discussed subsequently.)

! Separation due to changes in surface slope snd to impinging shockwa,_esfrom other surfaces can be alleviated by reducing the pressure

! gradient by dist_'Ibutiz_ the disturbance over some length. In other words,discrete shock wave_ are to be avoided. For instance, C_pman, Kuehn,

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