Wind-tunnel investigation of the effects of component build-up on the inlet performance of an 0.05-scale air induction model of the B-70 airplane at Mach numbers from 2.58 to 2.96

WATIOM^iL AERONAUTICS AND SPACE ADMINISTRATION

NASA MEMO 1-3-59A

for

UNITED STATES AIR FORCE

WIND-TUNKE]j INVESTIGATION OF THE EFFECTS OF COMPONENT

BUILD-UP ON THE INLET PERFORMANCE OF AN 0.05-SCALE

AIR INDUCTION MODEL OF THE B-70 AIR]?LANE AT

MACH NUMBERS FROM 2.58 TO 2.96

(COORD. NO. AF-AM-163)*

By Robert U. Hofstetter, Kenneth C. Endicott,
and Walter E. Isler

ABSTRACT

Exiierimental djtta are presented to show some effects of wing,
fuselage:, and canarcl on the inlet performance and external pressures
in the -vicinity of 1:he inlet. The comparative effects of several wing
configuiations on inlet performance are included.

INDEX HEADINGS

Air Inlets 1,4.1

Airplanes - Components in Combination 1,7.1.1

NASA MEMO 1-3-59A
for

UNITED STATES AIR FORCE

WIND-TUHTIEL INVESTIGATION OF THE EFFECTS OF COMPONENT
BUILD-IJP ON THE INLET PERFORMANCE OF AN 0.05-SGALE
AIR INDUCTION MODEL OF THE B-70 AIRPLANE AT
MACH NUMBERS FROM 2.58 TO 2.96
(COORD. NO. AF-AM-163)*

By Robert U. Hofs tetter^ Kenneth C. Endicott,
and Walter E. Isler

SUMMARY

An experimental investigation was made to determine the effects of
■vring^ fuselage^ and canard on the inlet performance and external pressures
of an D. 05- scale air induction model of the North American Aviation B-70
airplane .

T'Bst observations showed that the addition of the canard increased
the maicimum inlet pressure recovery by 1.1 percent over that obtained in
the presence of the fuselage forebody less canard. Lower surface wing
pressures outboard of the 24-percent semispan location were xniaffected by
the addition of the canard.

No significant change in inlet performance was caused by the addition
of both the fuselage forebody and canard.

Tvro of the wing configurations (North American Aviation Numbers I72
and 154 ) had a favorable effect on inlet performance^ while the third
(Number 1^5) wing affected performance adversely. The Number I72 wing
generaJ.ly exhibited more positive lower surface pressures than the Ni^m-
bers l^iU and 1^5 wing configurations .

INTRODUCTION

Tt.e North American B-70 airplane is an all-supersonic intercontinental
bomber designed to cruise at a Mach number of 3 '00. The successful

9 oa

t) § o

operation of the aircraft at this speedy over the range and with the
loading set forth in reference 1, requires that specific values of engine
perforiaance be realized. To this end considerable developmental work has
been done by the contractor on an air induction system which will satisfy
stipulated engine requirements. The outgrowth of this development program
has been a twin duct system utilizing two-dimensional-type inlets, and
combining variable internal geometry with a boundary-layer bleed system.

At the request of the Air Research and Development Command of the
United States Air Force an investigation was conducted in the 8- by 7~foot
test section of the Ames Unitary Plan wind tunnel on an air Induction
model of the B-70 airplane incorporating the North American Number 172
inlet version. The purpose of the investigation was to evaluate the
effects of wing, fuselage, and canard on the inlet performance and on
pressures external to, but in the region of, the inlet. Evaluation of
wing ej'fect on inlet performance included a comparative investigation of
severa], wing confiigurations to determine that which would produce a pres-
sure f:.eld, behind the leading-edge shock wave, most favorable to the
attainrient of good inlet performance. The results of this investigation
are re]5orted hereiii.

SYMBOLS

b
c
M
mc

mo

P

Pt

q.

^00

inlet capture area, 0.0933 sq ft

wing span, ft

wj.ng chord, ft

M£,ch number

m£-ss-flow rate based on free-stream conditions eind inlet capture
area, slugs/sec

di.ct mass-flow rate at engine face station, slugs/sec

st.atic pressure, Ib/sq ft

total pressure, Ib/sq ft

d;>-namic pressure, Ib/sq ft

P-Poo

pi'essure coefficient.

loo

a ai.gle of attack of basic model reference axis, deg

P angle of sideslip of basic model reference axis^ deg

X longitudinal "body axis in the vertical plane of symmetry of the
complete twin-duct system (denoted as the model reference axis
in figure 1; positive rearward)

X distance measured along the X axis positive aft of model station
0.00^ in.

Y lateral axis perpendicular to the vertical plane of symmetry of
the compj.ete twin-duct system, positive as shown in figure 1

y distance measured along the Y axis, in.

Z vertical axis in the vertical plane of symmetry of the complete
twin-duct system, positive as shown in figure 1

z distance measured along the Z axis, in.

Subscripts

CO free-stream conditions

2 c.uct station corresponding to the engine face location

MODEL

I'he basic mocLel tested was an 0.05-scale model of the B-70 airplane
air induction system furnished by Worth American Aviation, Incorporated.
It consisted, as shown in figures 1 through 3^ of an inlet-afterbody
combination with ei variable-geometry two-dimensional duct and true-scale
contours on one side of the model center line. The opposite side of the
model was unducted and designed so that the flow conditions occurring on
the ducted side simulated those existing during twin-duct operation.
Additionally, this offside portion of the model provided exits for the
b3.eed flow originating at the porous ramp surfaces on the ducted side of
the model. Mass -flow variation was accomplished by a remotely controlled
plug as illustrated in figure 2. The model was sting-mounted from the
wind-tunnel model support strut in an inverted attitude (see fig. 3)«

The basic inl.et used corresponded to the Worth American Aviation
Number 172 version. Figures 2 through 7 show pertinent inlet geometry
for this version, and indicate over-all areas available for the
application of boundary -layer bleed control.

VJhen wecig* extens2®ns»vere«e]Kipioiyed»in coniimcticn" <H>i'tek a supporting
wing configuration (see fig. 8) modification to the external geometry of
the Number 172 inlet version was possible. The extent of the alterations
resulting from the use of the wedge extensions is indicated in figures 9
and 10. Interchangeable segments of the cowl side plate were also used
to vary the cowl lip angle (see fig. 11 ).

The boundary-layer bleed capabilities of the Number 172 inlet version
were determined by the porosity of the plates attached to the areas avail-
able for bleed control. Variation separately of the porosity of these
plates^ and the effective porous area., altered the bleed characteristics
of the inlet. Four boundary-layer bleed configurations were used and are
tabulated in table I. Since a comparison of the effects of porosity varia-
tion is desired rather than a detailed evaluation of bleed mass flow, the
diminution of the net porous plate area caused by the webbing supporting
the plates is not considered.

Provision was made for the addition to the basic model of a wing, a
wing and forebody, and a wing, forebody, and canard. These components
were also furnished by the contractor. Several wings were tested in
conjuncition with the basic model to establish which configuration resulted
in a p:ressure field at the inlet conducive to the best inlet performance.
The va:rious wings are shown in figure 12, and a photograph of one version
is presented in fi,gure 13-

F:-gures l^f and 15 show the configuration consisting of the addition
to the basic model of a wing plus forebody. Figures l6 and 17 show the
configuration composed of the basic model and a wing plus forebody and
canard.

A summary of all configurations tested is given in table II.

MODEL COWOEIENT DESIGNATION

The model component identifying symbols and version nxunber designa-
tions established by North American Aviation, Inc., have been retained,
except where noted, and are defined as follows:

By Number 172 fuselage version. Including a forebody and the afterbody
portion of the basic model (fig. .15)

cowl side plate with an 11 cowl lip angle

(Subscripts 1, 2, 3, and h designate the boundary-layer bleed
configuration (see table l) used on the side plate.)

'■■ ^' O CO

C ' cowl side plate vith an o cowl lip angle

(Subscrli)ts 1, 2, 3^ and k designate the boundary-layer bleed
conflgur8.tion (see table l) used on the side, plate. )

Ct Number 172 canard version (fig. l6)

I Number 172 inlet version (schematically shown in fig. k)

(The first ramp extends from model station -8.6o to -5.8O, and
the bleed gutter height between the inlet and wing is 0.100 in.)

I2 Number 1^5 inlet version

(This is a modification of the Number I72 inlet version obtained
by extending the first ramp forward to model station -17-^55 by
means of wedge extension number 1 (fig. lO), used in conjunction
with the Number 1^5 wing version (figs. 8 and 9)> )

14 same as Ig above except the bleed gutter height between wing and

inlet is increased to G.165 inch

15 same as 1^ above except wedge extension number 2 (fig. 10 ) is

used

Ig same as I2 above except wedge extension number 3 (fig* 10 ) is
used

R movable ram]? system of the inlet

(Subscripts 1, 2, 2>f a-nd k designate the boundary-layer bleed
configuration (see table I) used on the ramp system.)

S upper and lower cowl plates

(Subscripts 1, 2, 3^ and h designate the boundary-layer bleed
configuration (see table l) used on the plates.)

Wrp Number I72 wing version (fig. 12)

Wij DTumber 15^ wing version (fig. 12)

Wip Number 1^5 wing version (fig. 12)

TEST PROCEDURE

Tlie test was conducted in the 8- by 7~foo't test section of the
Ames Unitary Plan wind tunnel. Pitch runs were made through an angle-
of -attack range from 0.2° to 4.2° at 0° angle of sideslip. Yaw runs
were mJide at angles of sideslip of -2.2° to +2.1° at 0° angle of attack,
and at 1.0° angle of sideslip at 2.2° angle of attack. All data were
taken at a nominal total pressure of 28o8 poimds per square foot absolute.

^The sjrmbol used by the contractor for this side-plate configuration is C^

S:lnce the inlet was located behind the wing leading edge^ the Mach
niuriber just ahead of the inlet was less than the free-stream value because
of the influence of the wing leading-edge shock wave. To provide inlet
perfornance comparison between configurations involving no component
additions ahead of the inlet with those which included a wing, the Mach
mamber schedule listed below was used.

Free-streiam Mach number

Free- stream Mach number

used for basic model

used for basic model

with no components

with a wing version

ahead of the inlet

added

2.880

2.955

2.700

2.799

2.i^98

2.578

Pi'essures existing in the model duct and over the external model
surface.'S were indicated on multiple tube manometers , and photographically
recorded. Figures I8 through 21 show static -pressure orifice locations
on all model components .

The total pressure existing at the engine face was obtained by
averaging 26 total-pressure tubes located at the engine face station.

Tt.e duct mass-flow rate existing at the engine face was determined
by use of an ISA flow nozzle located downstream of the engine face sta-
tion. , The flow nozzle was calibrated against a standard ASME orifice
plate to obtain a correction factor to be applied to the nozzle mass -flow
measurements .

Precision of Data

The uncertainties estimated to exist in the test data are listed
below :

Parameter

Uncertainty

a

±0.20°

P

±0.20°

Ptoo

±3.5 Ib/sq ft

K.

±0.005

I^tg/Ptoo

±0.002

P/Ptco

±0.002

Ap/q^o

±0.008

mo/mc

±0.003

o • c • •

RESULTS

Results of the present investigation are shown in the form of graphs
and tal)les. Plots of inlet performance showing effects of alteration of
the ba!5ic model, as well as the effects of free-stream conditions^ model
attitude, and wing addition, are presented in figures 22 through 29. No
inlet jperformance data are given for configurations including the forebody,
wedge extension nu;^"bers 2 and 3^ alteration of basic bleed gutter height
between wing and iiolet, and the cowl configuration employing no boundary-
layer bleed control. This omission is the result of failure of model
instruiaentation affecting the mass-flow rate measurements obtained during
tests of the af oreinentioned configurations . Qualitative analysis of inlet
perfonaance, based on pressure recovery alone, is possible however, and is
discussed in the following section.

Pressure coefficient data are presented graphically and in tabiilar
form (nee figs. 30 through 35) showing representative effects of Mach
number variation, and component variation on the pressures over the lower
wing surfaces. Table III is an index to figures 22 through 35.

External cowl surface pressure ratios are included (tables V and VI )
to sho\r the variations resulting from the addition of a wing and forebody
to the basic model. These data were obtained during runs in which solid
plates were faired over the cowl bleed areas preventing outflow through
the co\rl surface.

Pressure ratios obtained within the duct are also presented in
tables VII through XV. The data selected for presentation in this report
have been chosen to illustrate the effects on the internal flow of varia-
tion oj' free-stream conditions, model attitude, and component variation.
Table j'.V is an index to tables V through XV.

Tiie geometry of the subsonic portion of the duct was suitable for
use in assessing only incremental changes of inlet performance since it
was known to provide nonuniform flow to the engine. Presentation of
engine face flow distortion values would, therefore, not be meaningful
and are not included as part of the results.

DISCUSSION

The following discussion does not include an analysis of all the
data contained herein, but merely summarizes the significant findings of
the im'estigation relating to the over-all effects of component addition
on inlet performance . - '

Tlie four inlet bleed configurations tested are compared in figure 22.
Bleed <:onfiguration k was used in subsequent evaluations of the incremental
effects of component addition or alteration. (The scope of the investiga-
tion d:ld not include developmental refinements of the bleed system, and
bleed configuration h is not to be construed as necessarily optimum. )

Variation of inlet performance caused by the addition of the various
wing versions to the basic model is shown in figure 26. For the model
attitude and stream conditions noted in the figure, the Wrp and W^p wings
had nesirly the same effect on perfonnance - favorable, except near optimum
operati.ng conditions where a loss of 2.7 percent in pressure recovery

occurreid. The addition of the
inlet e,dversely.

Wt wing to the basic model affected the

Kg significant difference in optimum inlet performance for the basic
model -with the Wrp wing as compared to the basic model plus the Wt„
wing wa.s noted for variations in Mach number and angle of attack (see
figs. 27 and 28).

Presentation of inlet performance data showing the effect of the.
canard and the foreibody was precluded by model instrumentation failure
affecting mass flow measurements. Qualitatively, however, these effects
were indicated by pressure recovery data observed during the test. In
comparative runs, where the only variable was the canard, it was found
that maximum inlet pressure recovery for the canard-on configuration was
1.1 peroent_ higher than in the case of the canard-off configuration.
Comparative runs between configurations with and without the forebody-
canard combination showed no significant change in peak inlet pressure
recovery.

For the stream conditions and model attitudes shown in figures 30,
31, and 32, the W^ wing had generally more positive lower surface
pressures outboard of the 19-percent b/2 spanwise station than the
Wy and Wip wings.

The; addition of forebody and canard to the Wi3iIR4C4S4 configuration
(fig. 3!)) increased the lower surface wing pressures slightly at the
50-percent b/2 and 4o-percent b/2 spanwise locations. Ko effect was
appareni; at the 30-percent b/2, 2i4-percent b/2, or 19-percent b/2
locations. The addition of the canard (fig. 3^ had no effect on
pressures outboard of the 24-percent b/2 location.

Ames Res.earch Center

National Aeronautics and Space Administration
Moffett Field, Calif., Nov, 3^ 1958

REFERENCE

1. Aiion.: Weapon System 110-A, AF33(600)-3l801. Rep. No. WA-5T-39^-l>
North American Aviation^ Inc.

• ••

• • • (?•• p o*« ce

' » ^ C C CO

• 1500 COS on

• • • ..* i.

TABLE I.- BI^EEI; COMFIGURATIONS

Configuration 1

Bleed plate
location

Third
ramp

Bleed plate luet porous

porosity,
percent

8

area, sq. in.

13'^

Cowl
side plate

Cowl top

and
■bottom
plates

3i

2i+.21

8

26.17

-L

Bleed plate geometry,
dimensions in Inches

l*-»V.te-

3.05

U-2.8O-J

Fourth j

8

8.54

ramp

;

Fifth

8

16.28

ramp

3.05

5>3h

3.05

Note: Unshaded areas indicate portions of bleed plates through which
bleed flow may occur. 'I 0*^

org
»a e

Configuration 2

Bleed plate ;EJ^ea plate ^^^ ^^^^^^

T

location J JSS"^ ^

area., sq. in . ;

Bleed plate geometry,
dimensions in inches

Third
ramp

8

* — 4A2 •

6.93

T

3-05

4.__,^

»e.85-

,08

2.30

-.78

.^0 l»2.80-*
J.

Fourth

8 •

ramp

. Fifth

A

ramp

Cowl

elde plate

3i

Cowl top

and

.8

bottom

plates

21.12

2.30

Note: Unshaded areas indicate portions of hleed plates through which
bleed flow may occur.

11<

•• oot oo
C

TABLE I.- BLEED COMFIGURATIONS ■ Continued

Configuration 3

Bleed plate I ^^® Jjg J^?:*® i Net porous !
location [ percent iarea, sq.ln. ;

Third
ramp

8

Fotirth
ramp

8

Fifth
ramp

8

I Cowl
iside plate

Cowl top

and

hottom

plates

8

9.03

5.16

12.38

20.52

23.69

Bleed plate geometry,
dimensions in inches

— 4.42

1-2.86 H i;

-2. 80.

Note: Unshsxied areas indicate portions of bleed plates through which
hleed flow may occur . -| «^

O ??

C

* * -** • i.

TABLE I.- BLEED CONFIGURATIONS- Concluded

Configuration 4

Bleed plate 3leed plate Ijjet porous j
: location ; leycljt*^' ;area,6<i to. I

Bleed plate geometry,
dimensions in inches

Third
ramp

3*

13.i^8

Same as
Configuration 1

~^-

Fourth
ramp

3i

8.5^

Same as
Configuration 1

Fifth
ramp

16.28

Same as
Config\iration 1

Cowl
side plate

3i

2k. 62

Same as
Configuration 1

Cowl top

and
bottom
plates

3h

26.17

Same as
Configuration 1

13<

TABLE II.- SUMMARY LISTING OF C0IJFI3UIIATI0NS TESTED

A

Configuration
designation

Basic model inlet details

Wing
version

Forebody
version

Canard
version

Tnlet
version

Covllip
a^e,deg

Bleed
conf ig .

Wing -inlet "bleed
gutter, ht., in.

Wedge ext.
config.

IRCS

172

11

Ifolieed

0.100

IRiCiSi

172

11

1

.100

IR2C2S2

■ 172

11

2

.100

IR3C3S3

172

11

3

,100

1

IR4C4S4

172

11

h

.100

*
1

IR4C'4S4

172

8

h

.100

1
i
1

^^IR4C4S4

172

U

h

.100

172

1

«

WT2IR4C4S4

172

A

h

.100

I5U

•
•

¥T3l2R4C4S4

U5

11

k

.100

1

li^5

- "-51

*

WrpBrlRCS

172

11

HbMeea

.100

172

172

ff

•

WtBtIR4C«4S4

172

8
u

h

.100

172

172

•

WrpBjCr|iIR4C4S4

172

k

.100

172

172

s-

Wr[iBrjiC(jiIR4C'4S4

172

8

k

.100

172

172

172

WT2I4R4C4S4

172

U

h

.165

15if

¥T3l5R4C4S4

11+5

11

h

.100

2

145

^^3leR4C4S4

li^5

11

h

.100

3

11+5

3 ,i ^ .5
• ; J) o o o

TABLE III . - INDEX TO GRAPHICAL DATA

Subject

Configuration

Figure

Pressure recovery^
bleed effect

IRiCiSi, IR2C2S2

^3*^3^3^ IR4C4S4

22

Pressure recovery,
Mach number effect

IR4C4S4

23

Pressure recovery,
covl lip angle effect

IR4C4S4
IR4C'4S4

2h

Pressure recovery,
attitude effect

IR4C4S4

25

Pressure recovery,
wing effect

IR4C4S4
Wr]iIR4C4S4
WT2IR4C4S4

WT3IR4C4S4

26

Pressure recovery,
angle -of -attack effect
at three Mach numbers

WrrIR4C4S4

27

Pressure recovery,
angle -of -attack effect
at three Mach numbers

WT2IR4C4S4

28

Pressure recovery,
angle -o:?- attack effect

WT3I2R4C4S4

29

Wing pressure coefficients,

comparison of wing types for

various angles of attack

WriiIR4C4S4

WT2IR4C4S4
WT3IR4C4S4

30, 31, 32

Wing pressure coefficients,
Mach number effect

WTIR4C4S4

33

Wing pressure coefficients,
canard effect

WrpBTCTlR4C ' 4S4
WtBtIR4C'4S4

3h

Wing pressure coefficients,
forebody and canard effect

WriiIR4C4S4
WTBTCr[iIR4C4S4

35

15<

TABLE IV.- IKDEX TO TABULATED DATA

SulDJect

Configuration

Table

Bleed configurations

All

I

List of configurations tested

All

II

External cowl pressure ratios

IRCS

V

External cowl pressure ratios

WtBtIRCS

VI

Internal duct pressure ratios,
shoving effect of
bleed configuration

IRiCiSi
IRgCgSg
IR3C3S3
11x40404

VII

Internal duct pressure ratios,
shoving effect of Mach nvunber

IR4C4S4

VIII

Internal duct pressure ratios,

showing effect of

wing adc^itions

IR4C4S4

WrpIR4C4S4

¥t^IR4C4S4

Wt^IR4C4S4
3

IX

Internal duct pressure ratios,

shoving effect of
variation of covl lip angle

IR4C4Q4
IR4C'4S4

X

Internal duct pressure ratios,

showing effect of

variations of model attitude

iii^C^o^

XI

Internal duct pressure ratios,

showing effect of

wing, forebody, and canard

XR4C4b4

XII

Internal duct pressure ratios,
showing effect of the canard

WrpBTCrpIR4C ' 4S4
Wr|TBriiIR4C'4S4

XIII

Internal duct pressure ratios,

showing effect of variation

in "bleed gutter height

WT2IR4C4S4
¥T2l4R4C4S4

XIV

Internal duct pressure ratios,

showing effect of

the wedge extensions

WT3I2R4C4S4
%3l5R4C4S4
WT3I6R4C4S4

XV

16<

« 9 « « « «

TABLE v.- EXTERNAL SURFACE -PRESSURE RATIOS FOR THE IRCS
COHFIGURATION AT OPTIMUM INLET PRESSURE RECOVERY

(a) H.

-2.955,

p=o°

p =

=2.799,

= 0°

(C) Hx,

= 2.578,

13 = 0°

P/Pt.,

p/

Ptoo

P/Ptoo

Orifice

0^0.2°

a^2.2°

a^l+.2°

a^2.2°

a^l+.2°

ot=0.2°

a^2.2°

a^l+.2°

number

36

0.091^

0.092

0.116

0.108

0.138

O.II+8

0.128

0.132

36A

.07^

.072

.095

.085

.097

.101+

.096

.093

36B

.053

.036

.058

.06i+

.055

.080

.078

.060

36c

.o4o

.038

.033

.039

.031

.050

.01+7

.035

37

.038

.037

.035

.0^1

.038

.058

.052

.01+8

38

.Oi+0

.03^

.030

.039

.037

.057

.056

.053

39

.039

.OifO

.oi+o

.01+5

.01+6

.061+

.062

.063

i^0

.026

.028

.032

.037

.oi+o

.052

.052

.055

hi

.022

.022

.021

.025

.026

.038

.037

.037

k2

.096

.100

.055

.097

.07^

.113

.116

.095

J+2A

.070

.072

.050

.066

.060

.088

.081

.077

i<-2B

.oi+5

.oh6

.oi+i

.052

.05^

.075

.067

.065

i^2C

.Oi+0

.01+2

.01+3

.0I+9

.052

.067

.063

.066

^3

.039

.oi+i

.01+2

.01+6

.01+9

.059

.058

.060

kk

.03^

.03^

.035

.039

.ol+l

.056

.054

.055

i^5

.038

.oii-i

.oi+o

.0I+7

.01+7

.066

.066

.067

h6

.030

.03^

.037

.0I+2

.01+5

.058

.057

.062

ki

.018

.021

.023

.026

.029

.039

.038

.01+2

hQ

.022

.029

.031+

.037

.01+2

.052

.053

.058

h9

.022

.027

.030

.03^

.038

.0I+8

.0I+8

.053

50

.028

.030

.033

.039

.oi+o

.055

.055

.057

51

.028

.030

.032

.037

.039

.052

.051

.051+

52

.027

.029

.033

.036

.039

.050

.050

.055

53

.oi+5

.0I+9

.0^5

.0I+8

.01+7

.067

.056

.063

^h

.032

.035

.Olj-0

.0I+5

.048

.061

.061

.066

55

.031

.035

.0I+3

.01+5

.060

.060

.065

^6

.031

.033

—

.Ol+l

.01+1+

.051

.055

.059

57

.028

.030

—

.037

.oi+o

.052

.052

.055

58

.029

.032

—

.037

.oi+o

.050

.050

.053

59

.030

.035

—

.oi+i

.052

.058

.060

.069

60

.030

.033

.Ol+l

.01+6

.057

.058

.063

61

.032

.035

.01+3

.01+7

.062

.063

.065

62

.029

.033

.ol+l

.01+1+

.055

.055

.060

63

.028

.030

.036

.ol+l

.025

.050

.053

61+

.028

.030

.036

— — — ^

.037

.0I+9

.oi+8

.050

r?<

TABLE VI.- EXTERNAL SURFACE -PRESSURE RATIOS FOR THE WqiBtIRCS
CONFIGURATION AT OPTIMUM IKLET PRESSURE RECOVERY

(a) Hx. = 2.955, P = 0°

(b) Ho = 2.578, p = 0°

p/pt^

P/Pt^

Orifice
number

a = 2.2°

a = 4.2°

a =2.2° a -4.2°

36

0.107

0.110

0.157

0.161

36A

.072

.084

.109

.129

36B

.067

.065

.117

.125

36c

.Okk

.050

.075

.082

37

.051

.054

.075

.079

38

.oi+o

.045

.061

.068

39

.oi<-8

.038

.060

.067

ko

.032

.035

.060

.063

41

.070

.057

.045

.050

k2

.085

.103

.133

.120

k2k

■073

.076

.097

.096

J+2B

.060

.052

.083

.089

i+2C

.048

.050

.065

.072

^3

.0^3

.052

.067

.077

kk

• OifO

.044

.063

.063

i^5

.035

• 039

.060

.069

k6

.031

.035

.055

.061

hi

.023

.027

.04l

.046

k8

.029

.034

.050

.057

k9

.029

.032

.048

.054

50

.033

.037

.055

.063

51

.029

.033

.050

.055

52

.030

.034

.050

.056

53

.046

.053

.065

■ 079

^h

.037

.o4o

.065

.070

55

.035

.o4o

.067

.073

56

.035

.038

.056

.063

57

.030

.033

.052

.057

58

.030

.033

.049

.055

59

.033

.036

.057

.064

60

.032

.035

.057

.062

61

.037

.039

.069

.075

62

.034

.038

.057

.062

63

.030

.033

.051

.056

6h

.029

.031

.048

.053

18<

TABLE VII.- INTERNAL DUCT SURFACE-PKESSURE RATIOS AT OPTIMUM INLET PRESSURE RECOVERY FOR TEE BASIC
MODEL EMPLOYING FOUR DIFFERENT BOUNDARY-LAYER BLEED CONFIGURATIONS; K^ = 2.88, a = 0.2°;, p = 0°

I

S

Orifice
number

^/Pt^

Orifice
number

p/pt

IR3_CiS3_

XJa^ ^ 2 2

IRgC^^a

XR^C^Q^

IRiCiSi

-^■^2^2 2

'^^3^3^ 3

IR4C4S4

65

0.056

0.057

0.055

0.057

91

0.210

0.216

0.223

0.229

66

.071

.071

.069

.ori

92

.676

.378

.306

.420

67

.093

.095

.091

.077

93

• 364

.362

.369

.476

68

.096

.098

.094

.097

9^

.363

.509

.360

.488

69

.052

.053

.052

.053

95

.366

.595

.535

.602

70

.053

.055

.052

.055

96

.628

.614

.750

.729

71

.052

.054

.052

.054

97

.615

.606

.600

.663

72

.068

.068

.067

.069

98

.616

.591

.600

.653

73

.069

.070

.069

.070

99

.648

.603

.612

.670

7^

.095

.091

.095

.088

100

.680

.625

.643

.717

75

.121

.123

.135

.132

101

.646

.573

.612

.651

76

.267

.222

.243

.273

102

.834

.731

.795

.820

77

.221

.173

.239

.346

103

.856

.744

.822

.845

78

.308

.293

.322

.269

104

.864

.731

.830

.854

79

.280

.425

.321

.421

105

.869

.729

.834

.858

80

.500

.533

.543

.496

106

.870

.728

.835

.859

81

.625

.570

.595

■.582

107

.244

.303

.253

.167

82

.653

.579

.632

.619

108

.121

.168

.123

.220

83

.795

.663

.747

.771

109

.11?

.235

.123

.188

.84

.831

.728

.792

.693

110

.074

.211

.073

.085

85

.853

.761

.820

.844

111

.873

.784

.837

.863

86

.864

.773

.830

.854

112

.287

.300

.295

.073

87

.869

.777

.833

.857

113

.044

.108

.042

.046

88

.873

.784

.837

.864

114

.097

.255

.081

.082

89

.137

.135

.134

.133

115

.203

.449

.169

.126

90

.167

.165

.166

.163

116

.871

.780

• 835

.860

TABLE VIII.- INTERNAL DUCT SURFACE -PRESSURE RATIOS AT OPTIMUM INLET PRESSURE RECOVERY FOR THE BASIC

MODEL AT FOUR DIFFERENT MACH NUMBERS; a = 0.2°, (3 = 0°

IR4C4S4 configuration

Orifice

number

P/Pteo

Orifice
number

^/Ptoo

M=o = 2.^98

M)o = 2.700

Mx =2.880

M=o= 2.955

Moo= 2.498

H>o = 2.700

M«,= 2.880

Mx. = 2.955

65

0.103

0.07^

0.057

0.050

91

0.271

0.223

0.229

0.213

66

.^^6

.088

.071

.063

92

.330

.341

.420

.411

67

.126

■ .108

.077

.085

93

.258

.287

.476

.389

68

.125

.109

.097

.089

94

.276

.373

.488

.422

69

.091

.068

^053

.048

95

.278

.512

.602

.391

70

.092

.069

.055

.048

96

.639

.661

.729

.501

71

.091

.068

.05*+

.048

97

.672

.634

.663

.486

72

.113

.086

.069

.061

98

.683

.642

.653

.415

<iMi

73

.118

.088

.070

.063

99

.701

.662

.670

.394

■

7^

.113

.098

.088

.081

100

.739

.710

.717

.636

■

75

.076

.061

.132 '

.045

101

.681

.650

.651

.613

■

76

.086

.158

•273

.254

102

.815 ,

.815

.820

.796

^

77

.136

.174

• 3^6

.076

103

.835

.838

.845

.824

fO

78

.318

.304

.269

.252

104

.846

.849

.854

.832

o

79

.335

.286

.421

.376

105

.855

.856

.858

.837

A

80

.269

.i+30

M6

.364

106

.859

.860

.859

.839

81

.617

.58U

.582

.371

107

.265

.251

.167

.139

82

.662

.630

.619

.573

108

.325

.336

.220

.151

83

.772

.768

•771

.739

109

.533

.521

.188

.118

81^

.810

.810

.693

.791

no

.620

.589

.085

.061

85

.827

.835

.844

.824

ni

.862

.863

.863

.842

86

.8ii5

.850

.854

.833

112

.277

.314

.073

.043

87

.856

.856

.857

.836

113

.308

.362

.046

.038

88

.865

.863

.864'

.842

n4

.550

.507

■ .082

.038

89

.134

• 133

.133

.129

115

.606

.595

.126

.055

90

.166

.174

.163

.155

116

.859

.860

.860

.839

• ; 3 5

• JO.) .a

* * «

I

©23!

* 1? -> 4

*0 3<|

TABLE IX.- IWTEEWAL DUCT SUFtFACE-PKESSURE RATIOS AT OPTIMUM IKLET PRESSURE RECOVERY FOR THE BASIC

MODEL AT M^ = 2.88, AM) THE BASIC MODEL PLUS WING ADDITIONS AT N^ = 2.955; a

0.2°,

p = 0'

Orifice
number

P/Pt^

Orifice
number

^/Pt^

XJA4.C/4.O4

WTrR4C4S4

WT2IR4C4S4

WT3IR4C4S4

IR4C4S4

Wr[iIR4C4S4

WT2IR4C4S4

WT3IR4C4S4

65

0.057

o.oi+9

0.051

0.056

91

0.229

0.194

0.195

0.P2P

66

.071

,062

.063

92

.1+20

.1+16

.4o7

67

• 077

.o8ii-

.088

.089

93

.i+76

.396

.389

.392

68

.097

.772

.kih

M^

9h

.1+88

.1+12

.1+21

.420

69

.053

.okk

.oi+8

.061+

95

.602

.539

.i+36

.533

70

.055

.oh6

.050

.061+

96

.729

.707

.69k

.711

71

.054

.oh6

.01+9

.061+

91

.663

.705

.657

.603

72

.069

.060

.063

.061+

98

.653

.611+

.61+0

.587

73

.070

.062

.065

.061+

99

.670

.637

.660

.610

1^

.088

.770

.083

.076

100

»717

.680

.707

.621

■

75

.132

,782

.01+8

.061+

101

.651

.627

.61+5

.595

■

76

.273

.263

.259

.279

102

.820 ■

.798

.696

.790

■

77

.3^6

.308

.220

.308

103

.8^5

.836

.835

.815

w

78

.269

.236

.256

.21+1+

IOI+

.851+

.837

.81+3

.823

^ &

79

.ij-21

.377

.21+6

.379

105

.858

.843

.81+8

.826

^5

80

.h96

M^

.501

.k69

106

.859

.81+9

.851

.831

81

.582

.556

.601

.533

107

.167

.11+6

.134

.089

82

.619

.609

.630

.578

108

.220

.731

.156

.114

83

.771

.750

.763

.739

109

.188

.782

.126

.024

81+

.693

.794

.807

.668

TIO

.085

.066

.078

,068

85

.dkk

.822

.833

.816

m

.863

.81+9

.853

.834

86

.8^k

.838

.81+5

.823

11?

.073

.038

.158

.143

87

.857

.81+5

.81+8

.825

113

.01+6

.038

.045

.039

88

.86i|-

.851

.853

.833

11I+

.082

.037

.150

.122

89

.133

.11+8

.138

.126

115

.126

.053

.156

,122

90

.163

.165

.161

.Ikk

116

.860

.81+6

.850

.831

t 3 :j o
I -0 .; J o

TAEiLE X.- INTEMAL DUCT SURFACE -PRESSIHE RA.TIOS AT OPTIMUM INLET
PRESSURE RECOVERY FOR THE BASIC MODEL EMPLOYING CONSTANT
INTERNAL DUCT GEOMETRY AND VARYING COWL LIP ANGLES; N^=2.88o,
a = 0.2°, p =0°

Orifice

P/Ptoo

Orifice

P/

^t^

number

nuniber

Jj\^O^iD^

IR4C'4S4

IR4C4S4

^40*484

65

0.057

0.055

91

0.229

0.289

ee

.071

.070

92

.1+20

.330

67

.077

.092

93

.1+76

.171+

68

.097

.095

9k

.1+88

.1+15

69

.053

.052

95

.602

.1+1+3

70

.055

.103

96

.729

.575

71

.05^

.103

97

.663

.623

72

.069

.067

98

.653

.651+

73

.070

.070

99

.670

.671+

Ik

.088

.087

100

.717

.689

75

.132

.01+9

101

.651

.711

76

.273

.313

102

.820

.817

77

.3^6

.175

103

.81+5

.81+1+

78

.269

.232

101+

.851+

.857

79

.i^2l

.330

105

.858

.857

80

.I196

.501+

106

.859

.860

81

.582

.590

107

.167

.168

82

.619

.631+

108

.220

.226

83

.771

.769

109

.188

.208

Qk

.693

.811+

110

.085

.100

85

.81+1+

.81+2

111

.863

.862

86

.851+

.853

112

.073

.056

87

.857

.857

113

.0I+6

.ohk

88

.861+

.863

Tll+

.082

.039

89

.133

.16.1+

115

.126

.115

90

.163

.ll+O

116

.860

.860

2Z<

TA:5LE XI.- INTERNAL DUCT SURFACE -PRESSURE RATIOS AT OPTIMUM INLET
PR13SSURE REC0V:ERY for the BASIC MODEL AT THREE ATTITUDES; M^=2.880

IR

4C4S4 configuration

P/Pt

p/p+

Orifr.ce
number

^00

Orifice
number

^00

a=0.2°

a = 2.2°

a- 0.2°

a=0.2°

a = 2.2°

a = 0.2°

p=o°

(3 = 0°

P = 2.2°

p = o°

p = o°

p=2.2°

65

0.057

0.050

0.066

91

0.229

0.212

o.?n

66

.071

.06k

.080

92

.if 20

.430

.337

67

.077

.089

.101

93

.hie

.396

.318

68

.097

.097

.105

9k

.488

.452

.^39

69

.053

.052

.062

95

.602

.570

.428

70

.055

.053

.063

96

.729

.700

.451

71

.05^

.0i^9

.062

91

.663

.622

.509

72

.069

.065

.078

98

.653

.623

.632

73

.070

.066

.080

99

.670

.6ij-2

.631

Ih

.088

.088

.091

100

.717

.692

.693

75

.132

.oh9

.0i^8

101

.651

.620

.646

76

.273

.276

.243

102

.820

.800

.8o4

77

.3^6

.336

.192

103

.8if5

.830

.831

78

.269

.252

.286

io4

.Q^k

.8i^-3

.842

79

.U21

Mh

.256

105

.858

.850

.849

60

.496

.^^70

.273

106

.859

.856

.657

81

.582

.5^3

.579

107

.167

.193

.310.

82

.619

.599

.609

108

.220

.209

.304

83

.771

.7^5

.756

109

.188

0I75

.361

m

.693

.796

.800

110

.085

.078

.267

85

.Qkk

.825

.826

111

.863

.858

.854

86

.Q^h

.8I13

.8i^2

n?

.073

.093

.330

87

.857

.851

.8i^9

113

.0h6

.okQ

.252

88

.861^-

.858

.856

iii^

.082

.102

.444

89

.133

.137

.133

115

.126

.141

.532

90

.163

.161^

.173

116

.860

.855

.852

23<

n 5 g

B1> 9

TAI5LE XII.- IN'EERMAL DUCT SURFACE -PRESSURE RATIOS AT OPTIMUM IKLET
PRESSURE RECOVERY FOR THE BASIC MODEL AT IV^ = 2.880, AND THE
BASIC MODEL PLUS AWING, FOREBODY, AND CANARD AT M^„ = 2-955;
a = 2.2°, p = 0°

Orifice
number

pM.,

Orifice
number

^/^t^

IR4C4S4

WrpBrpCrpIR4C4S4

XR^O^b^

¥rpBrpC5.IR4C4S4

65

0.050

0.055

91

0.212

0.209

66

.061^

.072

92

.430

.560

61

.089

.093

93

.396

.527

68

.097

.099

9^

.ii52

.528

69

.052

.05^

95

.570

.528

70

.053

.056

96

.700

.585

71

.oii-9

.059

91

.622

.508

72

.065

.072

98

.623

M9

73

.066

.076

99

.61^2

M9

7^

.088

.091

100

.692

M6

75

.oJ+9

.0i^6

101

. .620

.333

76

.276

.295

102

.800

.768

77

.336

.317

103

.830

.798

78

.252

.h6-i

lOi^

.8i^3

.809

79

Mh

M'^

105

.850

.816

30

.ij-70

.375

106

.856

.821

31

.5^3

.407

107

.193

.174

32

.599

.338

108

.209

.117

33

.7^5

.IQk

109

.175

.099

Bij-

.796

.739

110

.078

,ohk

35

.825

.792

111

.858

.823

86

.8i<-3

.813

112

.093

.oi^3

37

.851

.819

113

.Oit8

.0^3

88

.858

.823

lli^

.102

.oif-3

89

.137

,li^9

115

.liH

.05*^

90

.164

.196

116

.855

.818

24<

TAB]J1 XIII.- IWZERWAL DUCT SURFACE-PEESSUEE RATIOS AT OPTIMUM INLET
].5RESSURE RECOVERY FOR THE BASIC MODEL WITH A WING, FOREBODY, MD
CANARD, AND WITH AWING MD FOREBODY; M^ = 2.955, a = 2.2°, p = 0°

Orifice

number

P/Pt.,

Orifice
nimber

I^/^t«,

WtBtCtIR^C'^S^

W-pBrpIR^C'^S^

WrjiBr]iCrrIR4C'4S4

Wr£BrrIR4C'4S4

s^

0.057

0.023

91

0.259

0.253

66

.072

.059

92

.320

.319

67

.092

.082

93

.370

.326

68

.098

.093

9h

.389

.353

69

.055

.050

95

.416

.426

70

.057

.051

96

.lj-71

.571

71

.060

.050

91

.58ij-

.595

72

.072

.062

98

.627

.630

73

.170

.203

99

.6h6

.649

Ih

.090

.083

100

.661

.662

75

.01^8

.051

101

.556

.536

76

.308

.297

102

.799

.796

77

.273

.165

103

.82ij-

.820

78

.293

.205

lOi^

.832

.805

79

.313

.312

105

.841

.837

80

.1^19

.i^95

106

.739

.698

81

.566

.577

107

.668

.667

82

.616

.619

108

.155

.164

83

.752

.7^9

109

.806

.173

81^

.77^^

.1^5

110

.078

.084

85

.816

.702

111

.848

.843

86

.820

.81^0

112

.043

.04i

87

.81^-1

.837

113

.047

.047

88

.81^9

.812

\\\

.051

.061

89

.173

.163

115

.102

.096

90

.153

.139

116

.844

.842

25<

TABLE XIV.- IIi(TEMAL DUCT SURFACE-PRESSURE RATIOS AT OPTIMUM INLET
PRESSURE RECOVEIRY FOR THE BASIC MODEL PLUS A WING CONFIGURATION,
USING TWO BLEED GUTTER HEIGHTS BETWEEN THE WING MD INLETj
^ = 2.955, a = 0.2°, p = 0°

Orifice
numoer

^/Pt^

Orifice
mamber

^/^t^

WT2IR4C4S4

%2l4R4C4S4

WT2IR4C4S4

WT2I4R4C4S4

6;?

0.053

0.051

91

0.261

0.195

6(5

.065

.063

92

.360

.023

61

.088

.088

93

.351

.389

68

.093

.411^-

9^

.516

.kZL

69

.043

.okQ

9^

.588

M6

70

.050

.050

96

.532

.69\

?:-

.Oi^9

.0J<-9

91

.601

.657

72

.063

.063

98

.621

.640

73

.065

.065

99

.651

.660

1^'r

.081

.083

100

.687

.707

T^

.138

.0l<.8

101

.557

.645

16

.268

.259

102

.700

.696

77

.320

.220

103

.673

.835

78

.259

.256

10i<-

.663

.843

19

.i^26

.21^6

105

.658

»848

80

.507

.501

106

.673

.851

81

.58i^

.601

107

.160

.134

82

.611.

.630

108

.200

.156

8-.

.754

.763

109

.099

.126

Qk

.781

.807

110

.150

.078

Ql'

.829

.833

111

.850

.853

86

,eh2

.8i<-5

112

.130

.158

87

.660

.8i^8

113

.0l<-2

.045

88

.850

.853

lll^

.121

.150

89

.139

.138

115

.139

.156

90

.li^2

.161

116

.8i^7

.850

26<

TABLE XV.- IlilTERWAL DUCT SUEFACE-HRESSURE RATIOS AT OPTIMUM lULET PRESSURE RECOVERY FOR THE BASIC
MODEL-PLUS -WING COHFIGURATIOIT EMPLOYING CONSTABT INTERNAL DUCT GEOMETRY AND THREE WEDGE

I

A

Tpvrnmvra TrvTT qtta-dcg .

'■'00

— o riKei

} y

Orifice
munber

P/Pt«,

Orifice

niuriber

^/Pt,,

WT3I2R4C4S4

WT3I5R4C4S4

%3l©R4C4S4

WT3I2R4C4S4

WT3I5R4C4S4

WT3I6R4C4S4

65

0.056

n n^^

c\ n^^

m

r\ 000

0.249

0.275

\^ • •^y-'

V ©^^w

y-'-

\j ,i~i-i-

ee

.066

.068

.067

92

.407

.455

.454

67

.089

.092

.091

93

.392

.378

.441

eQ

A35

.091

.088

9"^

.420

.459

.491

69

.o6it

.066

.066

95 .

.533

.437

.458

70

.061^

.067

.066

9e

.711

.382

.456

71

.061^

.066

.066

91

.603

.454

.447

72

.061^

.066

.066

98

.587

.427

.418

73

.06i^

.066

.066

99

.610

.550

.531

7it

.076

.083

.080

100

.621

.630

.627

75

.06I1-

.127

.136

101

.595

.518

.496

76

.279

.293

.289

102

.790

.732

.772

77

.308

.307

.339

103

.815

.801

.802

78

.2>^

.249

.294

104

.823

.815

.813

79

.379

.362

.380

105

.826

.822

.8?1

80

M9

.355

.369

106

.831

.793

.657

81

.533

.410

.413

107

.089

.137

.145

82

.578

.552

.545

108

.114

.155

.172

83

.739

.708

.709

109

.024

.093

.104

84

.668

.747

.747

no

.068

.123

.128

85

.816

.801

.801

m

.834

.829

.828

86

.823

.815

.800

IIP

.143

.069

.064

87

.825

.8?1

.818

113

.039

.037

.038

88

.833

.828

.828

114

.1??

.o4i

.038

89

.T?6

.\\h

.131

115

.122

.057

.082

•90

.Ikk

.166

.153

116

.831

.824

.823

Diotieiisloss In incbes
Ho scale

RefePMice axis

Model shovn in sn
inverted position

«p T O O -,>

«■: o .i

Relative viod

Figure i, - System of axes and positive direction of angles >

Model sta.
41.00

Mass -flow
coaatrol i^Lug

Sting sugpport

Side view

3-250

i

4.125

m

Z axis-

9

Note: Model sbo'wn stqoported

In an inverted position

Blmensloas in inches
No scale

Front view

Figure 2.- Side view of the "basic model incorporating the NtEaiber 172 inlet version.

Reference axis

Boarp BlsensiooasI Data

©3 3
• 3

Saoip
aibe

r leng^thj
In-

Baa^
heli^rt,

in.

Total rasBp
area,

aq, In-

Bebb^ «rea asrail- ^|
able for bleed ^M

8^ in- fl

1

2.821

3.250

9.168

0.000 3

2

1.681

3.250

5.^3

0.000 M

3

4.952 :

3.250

16.09J^

10.917 Ifl

It

2.85D

3.250

9.262

5.168 ■

5

13-750

varies

13.^55*

13.^95 I

6

8.1^^

tart.ieaj

20.065*

0.000 I

loBiaal Yalfies

1

]iotes- lSm&$ac9 an figure iztftlcste zaBp
deals>3&tl<m

Tbte aztgle iMitveen the first amS.
aecosA raa$ a:«rfaces sod tlie ref «
ersB&e asdU sqre fixed at 7* astd ,
11% reapectivaly.

Model is slKnut in imrerted ^c»t-
Sb seaOUs

Figure 4*> Scbooatle dravriag of tlBBft utaiber ITS inlet versicm, vitb \»ott<» ccfWI plate 3r«DdvtKL i^tiOving

tbe TGoedflible reaep attest.

Reference axis

A

Sbte- HvBBbers on figure indicate
rasrp designation

Typical grid area available for
the application of "boundary -layer
bleed control

Ko scale

Figure 5.'

Ifodel is shown In inverted
position

Schosatic side view oT tike vHariable raap systoi ^own in figure 3 showing relative locations!
of ranip JBz«as available for 'bom^tsaj-la.jei' bleed control.

»« • > I

Reference ajcls

Model sta.
-8.60

Model sta.
0.00

Area available for
bleed contjrol

All dimensions In Inches
No 3caJ.e

Note: Top cowl plate ahowi

Bottoin cowl plate Identical

Figure 6, - Schematic plan -view detail of top and bottom cowl plates of the

Number 172 inlet version.

33<

O '^T ,T

3.000

Reference axis

Area available for
"bleed control

Model ata.
0.00

All dimensions in inches
Wo scale

Figure 7.. - Schematic side-view detail of cowl side plate.

34<

A-23851

Figure 8.- Modification of the inlet from the Kvtmber 172 version to the
Number 1^5 version hy^use of vedge extension 1 and the Wm ving
version. iiJ^S"^-

Lateral displacemsttt of reference axis
Typical for all vedge extension
modifications »1 . 2k inches

Wedge extension No. 1

0.16 inches - Typical for all
wedge extension modifications

TjiTplcal for all
vedge exi^sion

modifications

Model sta.
-IT. ^55

Model Sta«
•8.60

Cowl side
plate

Model sta.
0.00

Note; Bottom cowl plate not shown

No scale

Figwre 9.- Drawing of the Niaiber 172 inlet version modified by the use of a
typical wedge extension to the Nvmiher ik^ version.

36-

> » •

'• ••? <

Model sta.
-8.600

1.08

a^iraop

r^

III
III
Hi I
II I

nd

(a) Wedge extension Number 1 with 1 — and 2 — ramps

Isentropic ccanpression
surface

(b) Wedge extension Number 2

Same as Number 2 except
for larger top fairing .

(c) Wedge extension Nianber 3

Piigure 10.* Details of wedge extensions investigated.

37<

Section through
cowl side plate

-th

5 — rajnp

2— ramp

rd
3"^ — ramp

k — ramp

'— 1— reaii)

reap

(a) Cowl 0id« pXate with lip angle of 6' with mod©l i.

(b) Cowl side plate with lip angle of 11" wltJi model t,
Flgijre 11 »- Details of the cowl side plate showing the 8" and U* lip angle.

38<

Model sta.
-15.125

Model sta.
0.00

Model sta.
-8.60

Dimensions in inches
No scale

Reference axis

Wing data

Airfoil section forward of the k% element - NACA 66A002.5

Ai]-foil section aft of the k^ia element - h% to fOp. hexagonal, 205*^ thick

Aspect ratio - I.88O
Taper ratio -

?51eed gutter height between wing and inlet - 0.100 inches

(a) jlasic model with Number I72 inlet version and the W^ wing
Figu]'e 12.- Schematic drawing of basic model -plus -wing configurations tested.

39<

]i4odel sta.
-13.855

Model sta.
0.00

Model sta.
-8.60

Dimensions in inches
No scale

60^ seini8pan»19,50

Beferenee axis

Wing data

Airfoil section forward of the k^ia element - NACA 66AOO2.5

Airfoil section aft of the k3i) element - k3i to 70^ hexagonal, 2.5^ thick

Aspect ratio - I.880
Taper ratio -

Bleed gutter hel|(ht betwen ving and inlet - 0.100 inches

(b) Basic model with the Nvmher 172 inlet version and the Wrp^ wing

Figure 22,- Continued.

4©<

Model ata.
♦17.^55

Model sta.
0.00

Dimensions in inches
No scale

60^ 8ealBpan»19.50

Reference axis

> Wing data

Airfoil section forward of the U5«t element - NACA 66A0O2.5

Airfoil section a:ft of the k^lo element - k5i) to 70^ hexagonal, 2.5/& thick

Aspect ratio - 1.880
Taper ratio -

Bleed gutter bei^t between ving and inlet - 0.100 inches

(c) Basic mocilel with the Number 1^5 inlet version and the Wip- wing

Figure 12.- Concluded.

41-

A-23843

Figure I3.- Installation photograph afi^;W3.e Wqira.C^S. model configuration.

m

e n fl

<C

Model sta

Model sta.
0,00
Wing

Bl*«d gutter helfifkit » .33I

Porelsody

Detail of bleed gutter between
f orebody e^id wing

Note; Plmenslons In Inches
No scale

?igure lU.- Typical jnodcl configuration including the forebody.

43<

^

^

^

Figure 15.- Installation photograph of the WrpBrpIR^C ' 4S4 model configuration

Model $ta.
-55.575

10.60

Model sta. Model ata.
^16M 0.00

Coaard 25^ c location

Z

Model sta.
-25.96

Detail Bhovlng angular
incidence of canard

Canard data

Sweep of 25^ element • kj*

Airfoil section - NACA 66AOO3

Aspect ratio - 2.078

Taper ratio - .1591

Figure 16. - Typical model configuration Including the foretody and canard.

45-

"■^©NFIBSNflAt.

O «

• • e

■B, afartanide to tislllng idj^

For all wing versions
A»19,50 in.
B-6T.it-5 in.

Airfoil seetiona at all stations are:
forvaard of i^5^ elwdent, BACA 66A002.5,
aft of 45^ elttaent, U5% to 70^6 liexagonal,
2.5^ i^iek, except for lftadlng(-edge ex>
tension, vhich la a flat plate on lover
atarfaee.

t) © (D ®

Ho scale

NuDbera on figure are
orifice nvmlaera

Orifice

No,

Location

Orifice
So.

Location

Orifiee
So.

Location

li''/^^

^e

U>>^]

i>o

$0

1

50

5

13

ItO

85

25

19

5

12

50

10

Ik

30

5

26

19

10

3

50

20

15

30

10

27

19

20

''4-

50

35

16

30

20

28

ll^.70

3.54 ^

'5

50

50

17

30

35

29'

13^0

3*28

i5

50

65

18

30

50

30

12.12

2.96

7

ko

5

19

30

65

31

10.83

2.65

3

ko

10

20

30

85

32

9.51^

2.36

!?

ko

20 ■

SI

2k

5

33

U.60

2,8if

10

ko..

35

22

2k

10

3^^

0.00

3.13

i:l

kQ

50

23

2k

20

35

7.00

5.06

111

ko

65

2k

2k

35

Figux« 18, • Location of preasure orlflcea on tbe lo^rer surface of all ving veraiona.

00
A

.80—

21.90

12.00

- l^.82-~^

Model sta.
0,00

^ ^ ^

Numbers on figure
are orifice nioDbers

-38-98

-30.92-

-Reference
axis

^ W^

2.22

.i

^~n

J^.125

o -J o c

o

0000

<9 a

^ I 6.025 r'^

Dimensions In inches

No scale

« o O 3 4

Figure 19.- Locaticaa of exteamal pressure orifices on bottom cowl plate.

■ju.yii

@ @€>

k.Q2^

Model sta*
0.00

21.90

-Reference axis

^^

f-^^

^^

-I — IJI3 typ.

-12.00

^^'^^^^^^ &=^

Numbers cm figure
are orifice numbers

DiBieBslons In Inches
Ko scale

Figure 20.- Locatlcsn of external pressiore oririces on co^rl side plate.

Reference exla -

Dimensions in inches
No scale

1. X -dimensions for raap surface
orifices are "based on ramps
positioned such that angles
betveen the raop surfaces and
the reference axis are;

7* for first ramp
11" for second reunp
l6" for third ramp
21* for fourth romp

a. Orifices numher 122, 013, ^^f
115 » and ll6 are not shovn, but
lie on the bottom cowl plate
directly opposite orifices 107,
108, 109, no, and Ul.

3. Model la shown in an inverted
position.

Orifice location

71^

3e,ln.

2*3

'».3

50

5.8

6.9

15.6

2, in.

Orifice
no.

Orinee location

x,ls.

z,ln.

85

I8.e

—19

86

21.1

-.3lt

87

21A

-53

88

26.9

-.725

89

0.8

90

2.0

91

3.5

92

5.0

93

5-3

9k

5.6

9595

5-9_^

_2I_

T.*

18.3

23-''

26.9

Orifice
no.

Orifice

107

li,6

li.08

108

5.9

U.27

109

6.6

U.35

110

7.3

k.k5

111

26.9

V.75

112

lt.6

1..08

h t-J

Ull

6.6

I..35

115

7.3

l*.U5

U6

26.9

lt.75

• 3 5 O O

• o o

♦ o c c o
e <£ 4

•■•«*-■

Figure 21.- Schematic drawing shoving locations of the duct internal static -pressiore orifices.

&

I

V

to
m

.u

. . . . 1 . . .
- ' • ' 1 • • '- •

;:;:;■;■'

-r— T

.;:;!.■.;
.:„|..._.

; ■ i - ; •
< ■

i ■ ' ■

- ; ■

;:■ :::•;

IReCsSe
D IRiCiSi
OIR3C3S3
^ IR4C4S4

.... _.. . . ...

.... J .......

T"":"

t- :■■:;:::■-

.— .-"^— .*.—

— "-^— • —

r^

:::.:l;.lu
: : : : 1 ; ■ : :

-:•--

.9

... i ... .

.:: :.G)

li

■<^^^2^^

.:' :k.::

;;;;;;■;

-::~"~r:-

I

.:....

........

--—

;:::•:;•:

-::\[\\:l:

V ■ 1 ■ •

::-:i:;-:

. . , ,

'.'■'■',

! 1 ; 1

....(..

•VijA;

m

..,.:....

. . . i . . . .

:::: ::::

• ■

^;;;;i:;i

';::i":-

;.;;;!•;;

i

fi

l:::\^

;:::li;;;

;;;•;

;:;'::;:

:;::

.:.;

:.■:

; ; ; • ■

■••;:;;;;

..,,.....

"*Ti~~"T

rtT T'p r. .

t:7"'-".'.'

-;;~H

1

>\\\\^

"TrrtrTTr

-; •~z*'-r^-. i.

TT! T

''. '~

*-••-! p*

-:rr.''-TT

■;r-~'-ff

','.'.'.]

■f^

Xi^-rrH--

■:•: ;-;;

•—-;-

•T

;;::;;;:;

'.'..'. '.'.'. 't

:;:::;::;

.J ..-'.,.. .

;;:::j/

ii

;-:^^^l!;^'^;-

'.'.'.'.]'.'•'.'.

-— r

;'';i^;"

:'•!:;:!':

■ ■ ■ -1 : : •

^;-;;"

'y^.'V:'

:';:i;:;^

:■;:::;■:

. . 1 . . :

rUf^-rrfrHf-i^r-:

T-H-r-r—

■.:■:■. .:■.■.
... t ... -

:;;:;::■.:

• b

::: j;r:

-r'-'r/r.-.

Tr:T-::rr^:T;-:-:--:r

;::;;;;;■■

\\\-M':\':

;:':i:'::p-::^::;;

:'.'.'.;'.'.'.'.

.:...,....

• ;;;;:: :;:':: i :': .:

•::;i::::r-;':::::

:::; 1 ;.;: 1 :;::;:;. :

i^^naaiiii^^^iiL

T;*7::':"'

-'"'-':-■'■

; : 1!: : : :

r

J : ; ,'r:,: ,■ ;

;;i;if;;i

••;::'

::::;:::!

;;:;

* . . . .

' ■ -

.7 .8

Net mass -flow ratlo> vIqMq

.9

1.0

Figure 22.- Effect of toimdary -layer bleed on the optim-um performance of
the Ni^mber I72 inlet version, vfith an 11" cowl lip, at Jlj„"2.880,
31=0.2% |3=0".

51-

n " r>

1.0

.9

8
(U

.)....
..;)..::

:.;!;;;.

. : 1 ...

. . .« . ...

1 — , '. ..■
:;::t.:-;

.:.i...

: : : ; 1 ; . : ;

1

"' ; : J ::t"

M:^

":yt::v.-

.:„.

_•-.: .

O Me,=»2.880
D Meri-2.700
OMo*«2'^98
A l4r«2.955

::;t.; .
— •-

■ - i • ' ■ •

::::!■:::

.:.:.:_U;^..

.;. >:;::-:

'J. :;;:.. . .

:: >.

:- :-; ;■:--

:':;:TT;:

"::

:.;::;:;;

-;;:- ::: f^

! . . : 1 , . :

.:. :.-.f?

^

: : : : 1 : . : :
.... 1 ... .

-'■.'■—':•:

1 ..;,... .

. : : :

r- • ■

::.:|::::|

--.'•-■ ".':

: ; . . i : . . .

W

i-i ' i '

7 ! I .'

::-.|::::

:::*^,

. . .,...-

: ; '. \ : ' .

:'-.:x:'.'.:

^Jn

■Tr:-

■'\\\'"'--

H--

: : : :

i,:.;.L

;;:L-;li;.

■. ' :

;.■■:;!;;;;

:: : :.::

'. 1 '. ^ I ' , . i . - * " ' . .
■ ' *■'"""?""! |7 ' ITT . '

.:i: :]-;.:.:. 'i'-.u:,:.

: : y

■A

.- •■. . ...

. . . , , .

. . . . . . .

--.■::t~:-

'.''.'...'.' !.'■'"•".'"

■..x

■ ;:: •,■..;.■,;•■■. i ;:. :

: -.1 . .;••:;
i

'..'.'. I

■3

?'" ." i '" '7

L^ ' U. ' ' ' '

.::.;. i';-.::L^:.-i.iii-

. . . . . .

:■::!::::

::..i::::

' '. ! t I ! ^ ! ;

.......

. .. . i

■ ■ - • ! ■ ■ ■

' ■ . ■ - ■
■ : : • i : . :

,...(...-

;;:;i:;.;

:■•-■-. r':r:

— ,_.i

Ql

•:::i:::; ;;;:i:::'

j.'.li

• ■ . - J

■ ■ ■ : i : : : :

','.'.'.,'.

..'''•'.'.'.',

:;::';■;

1 .' i " i ! ' ! '

:.;.;..-.U:.;-

-: ...I'.: .:.

'. . ' .\ ' '.'.'

:M^^-

■ i • —

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Figijre 264 - Effect of wing addition to the basio model on optitDum inlet

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Figure 30. • ConclviflLed.

61^

Pi

I

o

»

Spanwise location - 50^ t>/2

.2

;ffr

• . ■ f

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::::

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1 . . '.' . ~T7

I . . i

~

1

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i ^ ' ' 1 ! ! ! 1
... i ... .

- ■•I'" ;

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:-f:-

O Wr|iIB4C4S4
D WTgIR4C4S4
OW^ IB4C4S4

ii:^

1

B

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nlmi^-:

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20 1^0 60 80 100

Percent vlng chord, ^c

, (a)

Pig' ore 31.- CompariBon of pressure coefficients on the lower surface of
•the Nxanber 172, l^k, and li^5 wing versions for optlfflum inlet pres-

uure recovery at Mad-2.955» a»2.2', p=0*.

9i

O

8

to

n -J ;t
• c« «

Spanvlse location - 30^ b/a .

O WQ1IR4C4S4
D WTgIR4C4S4

O WijjgIR4C4S4

.2

SpanvlsiG

> locatlor

I -

21*^

'V

'2

u;;

'MI

—it

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tfr-

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....

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l-rr'r

....

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nif^

tit

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1

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-~

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rt

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i

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ip

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........

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rrrr-

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,z

^W

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::::

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; : : :

-rrrr

: fir

::'::

:i^:.

1,

ini.

M

•■•• ■■■■-

I * ' !

20 40 -60

Percent ving chord, ^

Figure 31. - Continued.

80

100

63<^

Spaawise location - 19^ "5/2

0)

g
I

to

.2

'.2

; : : :

fi

1

:i:: :::'

;l:;.r--'

, ... 1 ... :

-.'-

....

I wT

rrr/

■ T ■ • ■
.........

: : : : j : : : i

rr-.

WrpIR4C434
D Wrp2lR4C434

:;.';:

-r-fl

^;

........

: . : ; 1 ; : : ;

~y

■ ■ • ■

r-fr:

••:::

WT3IR4G4S4

rr-T-rr-^

■ ; ; : . : :
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....

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..... p, .

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r-fi"

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'•■-(•■■

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m

iil"r

.;:■•;

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2C

)

h

^0

60

80

100

Percent wing chord, ^

-^ p/<3oo

^W2

^

%IR4C434

WTgIR4C4S4

WqigIR4C4S4

lil.70

3.54

.03l«-

,he6

.430

I3.i^0

3.28

.03!^

,132

.202

12.12

2.96

.03i^

.070

.146

10.83

2.65

.OU

.026

.108

9'5h

2.36

.012

.018

.108

7.00

5.06

.026

.106

.146

l*.6o

2.81^

.010

.026

.122

0.00

3.13

.006

.001^

.122

(c)
Figxire 31.- Concliiaed.

64<

• •*• ••

••• •• *»s

Spanvise location • 50^ "b/2

^

p<

I

O

.2

.2 t

1 i '!

^iji

rrrr
'■III

TTTT

Jill'

' r ' '

: i i t

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rfrf

OWrpIR^C^S*

;-4
1 !.: ;

•1'; -ii'''

III:

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ri'i

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I ' 1 «

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DW»PgIH4C4S4

Ul f-

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:iil^

Tpir

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■'I 'I i i

nli

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:'..r:

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1

Spanwise locatlcai - 1^0^ 12/2

20

80

100

ko 60

Percent wing chord, ^

(a)

Flgvu.-e 32, ♦ compiirlBon of pressure coefficients on the lower surface of
tlie Ntnaber 17'-; 15^* an*! 1^5 ving versions for optimum inlet pres-
sure recovery at N^»2.955> Cf>k.2'', ^=0*'.

65<

no o 3 ©

Pi

<

o

9

n

Spanwlse location - 30^ b/s

Spaawlae location - 2ki) 'b/s

20

kO 60

Percent wing chord, ^

Pigvire ^.- Continued,

100

<

t)

8

V

0)

" *** "Spwiwfse location * 1^ b72

1^0 60

Percent wing chord, ^

100

^t V2

'k

Ap/cko

W^IR4C4S4

WijgIR4C4S-4

WT3IR4C4S4

U.70

3'5h

.061

.554

.khQ

13.i*0

3.28

.061

.238

.2kk

12.12

2.96

.149

.160

.304

10.83

2.65

.055

.080

.126

9.54

2.36

.051

.OkQ

.126

7.00

5. 06

.013

.136

.156

1^.60

2.8i^

.Oii-7

.037

.IU2

0.00

3.13

.033

.028

.142

(c)
Pigtire 32,- Concluded.

67<

p<

to

Spanviae location - 50^ b/2i

OJW"2.955
OH„=-2.579

WrpIR4C4S4

configuration

m

Trrr

m

t-r^

.'. I I = :

cs

U:il

;;.L

m:

rrrt.

Spanwise location - iK)^ t/a

20 to • 60 80 100
Percent vlng chord, ^

(a)

'.figure 33. - Effect of Mach nisniber variation on the lower surface pressure
coefficients of the Number I72 wing version for optimum inlet pres-
sure recovery at o»2.2*, ^=0*.

68'

ft
<

i

e a a

•/•«^

SpauwlBe location - 30^ b/2

T--

t:J.

fit

1 ii;

Hi;

nii

f 1 •■+■

TTT

r • + '

1::{-

t i t 1

lit:

lUi

1 if-t ,

iH:

l\:\

ih;ii;;

iiirT'it

^

fill

::;!"

11:1

Ijit

!;;•

. . 1 i..

(•.,-!

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■i »4^

J... .

■i'l

.; : : I

i..i

■* '■- t

; ! ;^
"1'

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::::

^

; • 1 '
J.: fw

: '1 ■ ■

'■) *■ J

:::r :::;

;:;!

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^

i?F

P

^

^

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"i ' '.

?St

iv;r^

f.T!-

;J-| t ;

t- I •

•iii

7^1

:::!

• H^

' ' . '.

i

^^

^;

; ' ; -K 7 ' ;

:;;:

T:':t
* • - »

T ' '"■^

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1^! '

:?:":•

-• ; .' ;

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:::;

' r 1 1

r ! '^.

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^1!:

:;;;'

:"■ i T

.- . -i . -|

t t-'-i

■S:l

t-it:

M iiii

, . , . .. - .^

::::!: :;:

.;;■;

1 ! ! :.

:■::.

In

... 4-+

•;fl

).:.:.

.. . , i.

li I .'

iiiiit:^:.

T- ? i f ' ^ ^ i

.:T!T|:::r

't.'-', '..il ', .* ;

• ■"•:[;:::

iWi

* * " *

i;:.:
■ft:;:

Jiii

ii:l

+ t . t

fnniH=!:

n-u:

■:::ii;:I;v

•:;:-)r:.;.;

::;:;r^-r;:

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

\[]] i]i-

i ■ ' •

■ • '■'

•|H}

Spanwlse location - 2h'f> t>/2

TTTT

;::f;

i:;Lf;.r*;' ■;.;.;'.

T;

Dl^»2.T99
OMce-2.579

Wrj;.IR4C4S4

conflgui'ation

Uo 60

Percent ving chord, ^c

Flgiare 33.- Continued.

80

100

69<

Pi
<

8

to
u

40 60
Percent wing chord, ^

100

i>'rt>/2

'k

Motf»2.955

Meo=2.799

Moo-2.579

Ap/q«,

Ap/q«

^p/q<«

14.70

13.40

12.22

10.83

9.54

4.60

0.00

7.00

3'5h
3.28
2.96
2.65
2.36
2.84

3.13
5.06

.034
.034
.034
.014
.012
.010
.006
.026

.0394
.0394

.27^3
.0283
.0195
.0666
'.0109
.1643

.0595
.0564
.4546
.0284
.0284
.0574
.0158
.0699

(c)
Figure 33,- Concluded.

70<

to
u

• • * •

Spaiiwise adcAtloii. - ^5C^

Spaawise location - kOf^ "b/s

-.2

20 1^0 60 80 100

Percent vlng chosrd, ^c

(a)

:?igtire 3!^. - Effect of canard on the lower sxarface pressure coefficients,
of the Wufflber I72 wing version, for optimum inlet pressure recovery
at 1V2.955, 06=2.2% p«0'.

"SpaQviaft'lVrfatlOn*'-' 305S i/l

<

I
1-1
u

o
o

«
«

Spanwise location - 2^ b/2

20 kC 60

Percent -wing chord, ^

(b)
Figure 3U.- Continued.

100

72<

=^

p<

to

kO 60

Percent wing chord,

^c

i \i/2

^c

^p/Qoo

W,j,BpCTlR4C4' S4

WrpRpIR4G4'S4

14.70

13.40

12.12

10.83

9.54

7,00

4.60

0,00

3.28
2.96
2.65
2.36
5. 06
2.84
3.13

.004
.056
.014
.008
.008
,036
.008
.014

.160

.062

.062

.012

,008

-.032

-.056

-.010

(c)
Figure 34. - Concliided.

73-

p<

<3

i)

•H
O

s

g

o

i

CO

I

Spanwlae location • 50^ b/s

Sjamwlse location • IfO^ Ij/s

20

100

1)0 . • .60

Percent wing chord, ^
(a)

]''ig-ure 35.- iSffeot of forebody and canard on the lower surface coefficients
of the Number 172 -i^dng version for optimum inlet pressure recovery
at M^=2.955; cj^2.2% p=0'.

74<

<

a;
o

<S3

Spanwiae location • 30^ 10/2

Spanwise location - 2^ b/2

i^O 60
Percent vlng chord, 56c

Figure 35.- Continued,

100

^pa^wlskiloe^tiosx I' ;:t9f>lt^£

4
<

u

Percent wing chord, ^c

i>-b/2

*c

A p /q,oo

WfrIR4C4S4

W|iBrjjCrpIR4C4S4

14.70
13.^^0
12.12
10.83

9.5^
7.00
4.60
0.00

3.54
3.28
2.96
2.65
2.36
5.06
2.84

3.13

.034
.034
.034
.014
.012
.026
,010
.006

.220

.058

.024
.090
.038
.052
.020
.024

(c)

Figure 35.- Concluded,

100

76'