C.P. No. 1346

PROCUREMENT EXECUTIVE, MINISTRY OF DEFENCE

AERONAUTICAL RESEARCH COUNCIL

CURRENT PAPERS

Measurement of the Internal

Performance of a Rectangular Air-Intake

at a Mach Number of 0.9

C. 5. Brown and E. 1. Goldsmith

Aerodynamics Dept., R.A.E., Bedford

LONDON: HER MAJESTY’S STATIONERY OFFICE

1976

PRICE f/-80 NET

UDC 533.697.24 : 532.542.1 : 533.6.011.35

*CP No.1346 January, 1973

MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9

bY

c. s. Brown E. L. Goldsmith

SLJMMARY

A rectangular intake having variable geometry compression surfaces and

designed for supersonic operation has been tested at a Mach number of 0.9 and a Reynolds number based on intake entry height of approximately 106. Tests have been made on the intake in isolation and on the side of a fuselage.

Maximum mass flow is less than estimates based on geometric throat size.

The deficit corresponds to an effective reduction in throat area of about 3 per cent. Critical point pressure recovery is lower than predicted, but the diff- erence can be related to a 'turning loss' factor as was the case at sup&sonic speeds.

With the compression surfaces of the intake horizontal, performance was unaffected by angle of incidence within the range 0' to 10'. With the compres-

sion surfaces vertical however, there was some loss in maximum mass flow and pressure recovery at critical flow conditions at angles of incidence above about 4O. This was only partly alleviated by removal of the swept endwalls.

The presence of this particular fuselage imposed no obvious effect on the performance of the intake.

__--. - -.._ -~~ * Replaces RAE Technical Report 73020 - ARC 34702.

2

CONTENTS

I INTRODUCTION

2 DESCRIPTION OF THE TEST RIG

3 DETAILS OF THE MODEL

4 DETAILS OF THE TESTS 4.1 Test conditions

4.2 Instrumentation and data reduction

4.3 Accuracy

5 DISCUSSION OF RESULTS

5.1 Isolated intake

5.2 Intake mounted on a fuselage with compression surfaces horizontal

5.3 Intake mounted on the fuselage with compression surfaces vertical

9

5.4 Comparison between vertical intake and horizontal intake

5.5 Engine face distortion 6 CONCLUSIONS Notation References Illustrations Detachable abstract cards

9

9

10

11

12 Figures 1-25

Page

3

3

3

4

4 4 6 6 6

8

.

-

3

1 INTRODUCTION

An extensive programme of wind-tunnel tests has been carried out at

RAE, Bedford to investigate the internal performance of a particular rectangular

intake both in a uniform flow field and on a fuselage.

Details of the supersonic performance at Mach numbers between 1.6 and 2.5

are contained in Refs.1, 2 and 3. This Report presents the results from tests

made on the same intake at a Mach number of 0.9. Tests have been made on the

intake in isolation and also when installed on the side of a fuselage forebody

with the compression surfaces of the intake both horizontal and vertical. In

the case of the vertical intake the effect of removing the swept endwalls was

investigated.

2 DESCRIPTION OF THE TEST RIG

Figs.1 and 2 show the intake, duct and forebody assembled on the General

Intake Test Rig used in the 3ft wind tunnel at RAE, Bedford. This rig has been

described in Ref.4. It consists of a sting support, a calibrated mass flow

control and measuring unit, a hydraulic actuator system for moving the intake

compression surface ramps and an instrumented duct with interchangeable exit

plugs for controlling and measuring the intake bleed flow.

3 DETAILS OF THE MODEL

The intake is designed for supersonic operation and the model is that used

in Refs.1, 2 and 3. The ratio of height to width at the entry plane is 1.54 and

the geometry of the compression surface ramps, cowl and bleed are as shown in

Fig.3. The first compression surface has a fixed angle 61 of 10' and the shock

from the leading edge theoretically falls on the cowl lip at a free stream Mach

number of 2.43. The second compression surface is movable and is linked to the

rear ramp. In the configuration in which the intake is mounted on the fuselage

with its leading edge vertical, the two movable surfaces are connected to the

hydraulic actuator system on the intake test rig. However, when the intake is

assembled on the fuselage so that its leading edge is horizontal this arrangement

is not possible and the movable surfaces although still linked together in the

manner shown in Fig.3, are controlled by means of a manually operated lead screw.

The gap between the second and rear ramps forms a slot for bleeding the

boundary layer from the compression surfaces and this slot extends the whole

width of the intake. The bleed air is discharged through a duct into the free

stream. Bleed exit area is varied by means of interchangeable plugs. In

addition to zero bleed, tests were done at constant ratios of bleed exit area to

intake entry area of 0.05, 0.101 and 0.174.

4

Two different shapes of endwall were used in these tests. One was a full

swept endwall in which the leading edge coincides with the line joining the

leading edge of the intake to the cowl lip. The other is a minimal endwall

which is sufficient to contain the space under the second ramp at maximum 6 2'

but otherwise has a leading edge which is vertical at the cowl lip. Details of

the two shapes are shown in Fig.4.

The area distribution through the intake and duct for various values of

&2 is shown in Fig.5a while Fig.5b shows the area distribution through the

intake alone between the cowl lip and the rear hinge for those values of 62

used in the tests. The ratio of engine face cross-sectional area to maximum

intake capture area is 0.88 and the distance from the cowl lip to the engine

face is 9.89 times the intake height.

The nose and canopy only of the fuselage are represented. Fig.6 shows

details of the forebody including the relationship between the fuselage datum,

the intake datum and the nose cone centre line. The model is mounted in the

tunnel on the intake datum line so that c1 is the angle of incidence of the

intake relative to the wind-tunnel free stream.

4 DETAILS OF THE TESTS

4.1 Test conditions

The tests were all made in the 3ft wind tunnel at RAE, Bedford, utilising

the working section with top and bottom slotted walls. Tunnel conditions were

such as would give a free stream Mach number of 0.90 in the empty tunnel. The

effect on free stream Mach number of the intake test rig which has the rather

high blockage ratio of about 3 per cent is not known and no corrections for

wall constraint have been applied. The Reynolds number based on entry height

was approximately 106.

4.2 Instrumentation and data reduction

The standard mass-flow control and measuring unit used in these tests is

fitted with a cruciform rake having a total of 24 pitot tubes for measuring

total pressure at the engine face station. The tubes are disposed for area-

weighted averaging and the rake is rotatable to enable pressure surveys to be

made in greater detail'. Static pressure at the engine face is measured by using

four holes equally spaced around the circumference. Details of the character-

istics and calibration of this type of airflow meter are to be found in Ref.5.

The bleed duct contains twelve pitot tubes arranged in three rakes of four

tubes for measuring total pressure and three holes for measuring static pressure.

5

Pressure recovery is defined as:-

n pf pB 1 PO3

or c=nP, c P.

J (1)

1

where P co is the free stream total pressure and P. J

is the pitot pressure at

the jth tube in the rake at the engine face station in the case of Pf or in

the rake in the bleed duct in the case of PB; and n is in either case the

number of pressure points in the survey.

The ratio of engine face mass flow to maximum capture flow was computed

from:-

A a3 ( ) A, f

n

c KfAf 1 = nA M

e 00 (2)

f 1 + 0.2M2 5

co

where M. = J

local Mach number calculated from P. and pf; pf being the J

average static pressure at the engine face

Ae = maximum capture area

Af = area at the engine face

Kf = discharge coefficient obtained by calibration5.

Similarly bleed mass flow ratio was computed from:-

Am ( ) Ae B

= (3)

where P. J

now refers to pressures in the bleed duct and M. is the local Mach J

number calculated from P. J and pB9 the average static pressure in the bleed

duct. pb is the area of the bleed duct, in this case used without a discharge

coefficient.

Total mass flow ratio

(?jT = (tgf +(t$ l

(4)

The parameter used to define flow distortion at the engine face is DCbO,

defined as

DC6O = p60 - Pf)

9f (5)

where p60 is the mean total pressure in the worst 60' sector and qf is the

mean dynamic pressure at the engine face, derived from P f and pf'

4.3 Accuracy

Errors in the direct measurement of engine face and bleed duct total

pressure and therefore in pressure recoveries based on free stream total pressure

are thought to be small, not more than 0.1 per cent. Uncertainty associated with

the mean tunnel Mach number and more particularly with the calibration of the

airflow meter probably means that the error in mass flow ratio could be as much

as half a per cent.

l

5 DISCUSSION OF RESULTS

5.1 Isolated intake

Figs.7, 8 and 9 show the internal performance of the isolated intake at

three different values of bleed exit area, including the case of zero bleed, and

for a range of values of second ramp angle 62. Pitching the intake with its

compression surfaces horizontal had no effect on the pressure recovery - mass

flow characteristics within the range of incidence investigated.

Fig.10 summarises the variation with bleed exit area of maximum total and

maximum engine face mass flow, maximum pressure recovery and critical point

pressure recovery. The effect on performance of internal contraction in the

duct and of the location of the throat in the duct is well illustrated. At

values of second ramp angle 6 2 above about 5' where the minimum duct area is

in the region of the bleed slot, the bleed becomes inoperative at maximum mass

flow owing to the low internal pressure.

7

At values of 62 between 2O and So it is possible to obtain small amounts

of bleed flow but as bleed exit area is increased the intake soon chokes in the

region of the bleed slot. At values of A2 below 2' the minimum duct area is

at the rear hinge and bleed mass flow increases quite normally with increasing

bleed exit area. For all values of g2 maximum engine face mass flow is

practically independent of bleed exit area. For values of 62 above about 2'

the large bleed exit area (AB/Ae) = 0.10 results in a slight reduction in

maximum mass flow. This is not clearly understood.

The apparent minimum duct area has been derived from measured maximum mass

flow and is shown in Fig.:l, together with the geometric minimum duct area and

the inlet entry area at the cowl lip. The ratio of apparent to geometric minimum

duct area is approximately 0.97 except when the throat is at the rear hinge, in

which case the ratio falls to a minimum of 0.955.

Critical point pressure recovery shows a very slight increase with

increasing bleed exit area (Fig.10). This occurs with or without bleed flow

except at &2

equal to 0' where with a large bleed flow pressure recovery falls

off by about 2 per cent. Estimates of skin friction and interference losses

based on the work of Ref.6 indicate that at critical flow conditions there are

large additional losses. In reports of tests at supersonic speeds 192 these

additional losses have been ascribed to the turning of the flow in the duct

downstream of the cowl lip. In Fig.12 the losses in the present case are

plotted against w, the angle of turning between the front and rear movable

ramps. The correlation is again quite good, although the loss at 62 equal

to go seems to be rather high. Some explanation of this may be contained in

Fig.14. This shows the pressure distribution on the vertical centre line of the

duct obtained from a pitot rake near the rear hinge position as shown in Fig.3.

For each value of &2 and bleed exit area the pressure distribution is shown

for successive points on the pressure recovery-mass flow characteristic, pro-

gressing from a supercritical condition, through critical, to a subcritical

condition. At values of 62 of 5 0 and above, where the minimum duct area is

in the region of the bleed slot, the pressure distributions indicate that at

supercritical flow conditions the flow at the pitot rake is supersonic. Fig.13

shows a comparison of the pressure recovery indicated by the pitot tubes in the

middle of the duct when the intake is just supercritical, with the pressure

recovery calculated from the assumption that the Mach number at the rake is

derived from the area ratio At'Pk and Mt = 1.0.

8

The correspondence is quite good. This expansion over the rear ramp will

inevitably be followed by a normal shock system which in general will be prac- tically extinguished at the critical point. Most of the additional losses will be associated with separation occurring on the surface of the rear ramp. How- ever, at the relatively high 62 of go, followed by a rapidly expanding duct

downstream of the throat, there is probably still some significant shock loss at the critical flow condition.

5.2 Intake mounted on a fuselage with compression surfaces horizontal

Pressure recovery-mass flow characteristics for the intake when installed on the fuselage with its leading edge horizontal are shown in Fig.15. The data were obtained at a constant bleed unit area ($/A,) equal to 0.10 and for values of second ramp angle 62 equal to -loo, 0' and 5'. As was the case with the isolated intake varying the angle of incidence had no effect on these characteristics within the range investigated.

In Figs.lGa and 16b maximum total mass flow and maximum engine face mass flow are shown plotted against second wedge angle 62 for the intake alone and

for the intake on the fuselage. Estimates of mass flow based on intake internal geometry have been calculated and are included for comparison. The presence of the fuselage causes a reduction in maximum total mass flow of about 1.5 per cent.

The difference in approach Mach number between the isolated intake and the intake on the fuselage is not known; all mass flows have been referred to nominal free stream Mach number. However, the difference is hardly likely to be the cause of this reduction in total flow because there is no corresponding reduction in engine face mass face. The deficit is entirely in the bleed flow quantity and is more probably caused by a change in bleed exit conditions due to the presence of the fuselage.

Measured maximum engine face mass flow falls below the quantity estimated from the intake internal geometry because of the apparent throat size referred to in section 5.1. The difference between the measured maximum total mass flow and the estimate based on the inlet area at the cowl lip has one of two origins depending on the particular value of 62. For values of 62 above about 2' it is due largely to choking at the bleed inlet and consequent failure of the bleed to operate. At lower values of 62 it is due simply to insufficient bleed exit area.

The negligible effect of the fuselage flow field on intake pressure recovery is illustrated in Figs.16c and 16d which show the variation with d2

9

of both critical point and maximum pressure recovery for the isolated intake and

the intake on the fuselage. Estimates of skin friction and interference losses

based on Ref.6 are also shown on this figure to illustrate the large additional

pressure recovery losses referred to in section 5.1.

5.3 Intake mounted on the fuselage with compression surfaces vertical

Figs.17 to 22 show pressure recovery-mass flow characteristics for the intake when mounted on the fuselage with its leading edge vertical. The intake was fitted with either two swept endwalls or two unswept endwalls. Details of

the endwall shapes are shown in Fig.4. For each endwail configuration data

are presented for a constant second ramp angle equal to -loo, with bleed exit

areas MB/A, 1 of 0.05, 0.10 and 0.174, and a range of incidence of the intake-

fuselage assembly from zero to 12'.

In Fig.23 the variation of mass flow and pressure recovery with bleed exit

area is summarised for both endwall configurations. Removing the swept end- walls reduces the large loss in maximum pressure recovery which occurs at an

incidence of 12'.

5.4 Comparison between vertical intake and horizontal intake

The mass flows and pressure recoveries of the various configurations of

the intake when mounted on the fuselage are shown in Fig.24 for a second ramp angle of -10' and a bleed exit area (AB/Ae) equal to 0.10. The indications are that at low angles of incidence (01 = Co to 4') there is little difference between any of the configurations. However, the intake with its compression surfaces vertical is affected by incidences above 4'. In particular the configuration with two swept endwalls suffers a significant loss in maximum pressure recovery at an incidence of 12'.

5.5 Engine face distortion

Distortion at the engine face expressed as DC 60

is sho-m in Fig.25. The maximum subcritical value of DC 60 is also shown, This is rarely greater than the value at the critical point and in any case always occurs within 5 per cent

of maximum mass flow. At low angles of incidence, up to about 4', the value of

DC60 is in the region -0.25 to -0.35 for all configurations. However, at angles of incidence beyond 4' the intake with the vertical leading edge shows some deterioration in flow uniformity whereas with the Leading edge horizontal the uniformity is apparently unaffected.

10

6 CONCLUSIONS

Measurements have been made of the internal performance of a rectangular

intake having variable geometry and designed for supersonic operation. The

measurements were made at a Mach number of 0.9 and a Reynolds number based on

intake entry height of approximately 106. The intake has been tested through a

range of angle of incidence of 0' to 10' in isolation in a uniform flow field

and in the environment generated by a particular shape of fuselage.

Maximum mass flow is slightly less than would be estimated on the basis

of geometric throat size; the deficit corresponds to a reduction in throat size

of about 3 per cent.

Internal contraction, where it is located in the region of the bleed slot,

leads to a failure of bleed operation at maximum mass flow.

Critical point pressure recovery is well below that predicted after taking

into account duct losses. However, as was the case at supersonic speeds, these

additional losses correlate well with the turning angle between the front and

rear movable ramps.

With the compression surfaces of the intake horizontal there was no

deterioration in performance with increase in angle of incidence; nor was the

performance impaired by the presence of this particular fuselage.

With compression surfaces vertical, the intake on the fuselage, suffered

some loss in mass flow and pressure recovery at angles of incidence above 4 0

This was only partly alleviated by removing the swept endwalls.

11

% % AR Af 4r K

L M

P

pR q P X a

% &2 w

NOTATION

bleed exit area bleed duct cross-sectional area

cross-sectional area at the rake station near the rear hinge

engine face cross-sectional area cross-sectional area at station X

discharge coefficient distance from the cowl lip to the rear hinge Mach number

total pressure pitot pressure at the rake station near the rear hinge

dynamic pressure static pressure distance downstream of cowl lip angle of incidence of intake relative to free stream

angle between first compression surface and free stream ahead of intake angle between first and second compression surfaces

angle between front and rear movable ramps

Subscripts

C at critical flow conditions

B in the bleed e in the intake entry plane f at the engine face i at the cowl lip t at the section of duct minimum area T total, i.e. engine plus bleed

X at station X al in the free stream

12

No. Author -

1 C.S. Brown

E.L. Goldsmith

2 C.S. Brown

E.L. Goldsmith

3 C.S. Brown

E.L. Goldsmith

4 E.L. Goldsmith

5 I. McGregor

6 J. Seddon

REFERENCES

Title, etc.

Measurement of the internal performance of a rectangular

air intake with variable geometry. Part I.

ARC CP No.1243 (1971)

Idem. Part II. The effect of incidence.

ARC CP No.1292 (1971)

Measurement of the internal performance of a rectangular

air intake mounted on a fuselage at Mach numbers from

1.6 to 2.

ARC CP No.1291 (1972)

Variable geometry intakes at supersonic speeds. Some

techniques and some test results.

AGARD CP 34 (1968)

The characteristics and calibration of two types of

airflow metering device for investigating the performance

of model air intakes.

RAE Technical Report 71212 (ARC 33894) (1971)

Boundary layer interaction effects in intakes with

particular reference to those designed for dual subsonic

and supersonic performance.

ARC R & M 3565 (19663

\

Cod internal angle 1J”

/

Cowl external angle 2-149 in

b-689 in lb9*9mm

Fig.3 Geometry of compression surface ramps, cowl and bleed

9-017 in

102 mm

c

Lea&q edge

Swept endwafl Unswept endwall

E E cu . s

a 66-lmm

e

Leading edqe

Fig.4 Details of endwalls

2.2

2-o

1.6

l-6 A, Ai I.4

I I I I I I I

- - - . a-

0-I 03 o-4 o-5

X (enqme face) .

Fig.Sa Area distribution through intake

I I

o-5 x I-0 0

Fig. 5b Area distribution through intake between cowl lip and rear ramp hinge

I-0

o-95

Pf

P,

0.90

O-85

I-0

PfJ o-9 p, O-8

o-7

O-6

O-7

i I I I I

03 o-4 (

0-j o-4 C

.

I-O

o-95

Pf

P,

o-90

0.65 1

,

*a3 0

0.04

ii- t B o-03

O-02

0.01 0

1

Fig.8 Variation of pressure recovery with mass flow. Isolated in take

I.0

Pf

PO0 o-95

pa O-8 - PO0 0.7

r ___I___ ___^.___ .- .__-. - --.--._-- -1

O-6

o-5

06 o”- IO0 AB (> -b,

Fig.9 Variation

o-95

0*9c

0.08

0.06 .

of pressure re Isolated intake

. ,

045

bo 0 A,

6, 0’ 625” I

A, 0-I ( > Ae

O-90 1 I I 1 , .

Fig.10 Variation of mass flow and pressure recovery bleed exit area. Intake alone. cL=O*-IO*

0.75

O-70

0.65

O-60 At A,

053

O-50

o-45

O-40

Ceometrlc ------

Derived from measured mass flow:-

0 0.05 0.10 0 X A

I I I 1

0 2 4 6 8 6,” 10

Fig. II hct minimum area derived from measured mass flaw.

Isolated intake

. * h

0

./

p”

0’

/ I

/ .

/ I I I I

- - - I - . -a

0 o-2 o-4 O-6 0.6

Anqle between front and rear movable ramPs O(radians)

. \ I

O-96

o-94

0.92

o-90

Calculated assuminq sonic throat

Measured o

I I I I 2 4 6 8

Second ramp anqle 6;

Fig.12 Correlation between flow turning Fig. 13 Pitot pressure at rear hinge loss and angle between front and rear position. Critical flow conditions. movable ramps. Isolated intake. Zero bleed Isolated intake

61 9”

6, 7”

62 5”

bt 2’

62 0’

I I

1 0 K

A

Cod side

Fig.14 Pressure distribution on duct vertical centre line at rem hinge position

b L

L n

I- 0

0 g Q)

o-95

O-90

00

0

0.6 4 P,

o-4

O-2

1 I , 1

0.5 0.4

I.0

Pf

0 a,

0.95

Fig.15 Variation of pressure recovery with. mass Intake with swept endwalls on fuselage with

O-70

0.65 ho

0 A eT max

0~60

\ \

\ \

I \

\

\ \ \

\ \

\ I

\ \ \ \ \

\

\ \

\ \

7

a

+kin friction

O-65

O*bO

o-55

ox

Istimated from ntakc qeometry--- leasured :- ntake alone Q ,ntake on fuselage A

0 5 62’ IO b

rice loss Ref 6

I

5 6 d

D

Fig. 16 a-d Variation of pressure recovery and maximum mass flow with 6,. Intake alone II on fuselage with leading

edge horizontal (*‘I&) = o- IOI

L ” r

I4

o-95

o-9(1

0.85 I

I-0

o-95 P f PO0

0%

O-85 ii 0.j 0.4 O-6 O-7

0.05

0.3 o-4 O-5 AS O-6

0

O-7 A

ef

Fig. I7 variation gf pressure recovery with moss flow. Intake with swept endwalls on fuselage with LE vertical

I-0

o-9: Pf

0*9(

O-85

0.9

pB O-8 P,

o-7

0-b

o-5 I I I I

3 o-4 0.7

I-0

o-95

Pf

Ki

o-90

0.8518 0-J 0.4 0.7

O-05, 0.04

o*o’J

0.02

o-01 n

6, = - IO” A*

1 0 A, = 0.05 a zoos 4” 0 &=8” x cx=12” A

1 I “0-j 0.4

I I I O-6 0.7

Fig. 18 Variation of pressure recovery with mass flow. Intake with unswept endwalls on fuselage with L E vertical

. .

I-0

o-95 P t L

o-90

O-85

I-0

o-9

P, O-6 PO0 O-7

,- O-6

o-5 1 I I I 0-j 0.4 O-5 A, O-6 O-7

0 A Q T

v-

, -

I-O

o-95

Pf

P,

O-9(

O-8!

O-10

o-09

0.06

o-05

Fig. I9 Variation of pressure recovery with Intake with swept endwalls on fuselage with

o-95 !i

O-85 I

0.6

0.5

62 = - IO0

I I

3 o-4 0-5 A, O-6 O-7

0 A, 1

cx= 0" and 4' 0

I ot = 8" X

oc = 12” A

I I 5 o-4 o-5

Cl B ii-6 d-7 A, t

O-95 3

iii

O-PO

O-85 I

Fig.20 Variation of pressura recovery with mass flow. Intake with unswept endwoils on fuselage with C E vertical

. .

I-0

o-95 P f pop

o-90

0.85

74 A lb

I I I I I

09 O-4 A, O-5 ( \

O-6 0.7 7

o-9

P, O-8 p, o-7

O-6

@ I.0

0.95

Pf

L

0.90

0.85 I I I I

03 O-6 O-7

O-08 1 I I I I

03 O-5 O-6 O-7

Fig.2 I Variation of pressure recovery with mass flow. intake with swept bdwalls on fuselage with LE vertical

I-0

0.95

!k PCQ

o-90

O-85

I-0

o-95 P f pap

O-90

O-85 i I I 1 ,

03 Q-4 O-5 A, O-6 LA

O-7

62 = -10” *f!

0 A, = o-174

o-5 O-4

O-6

0.5 I 1 I I I 0.5 0.4 O-5 0 A, 0-b 0-7

A, T

O-001 I 1 I I 0-J O-4 O-5 A, O-6

0

O-7

AC f

Fig.22 Variation of pressure recovery with moss flow. Intake with unswept endwalls on fuselage with LE vertical

O-65

0.60

d o’and 4”

4 A 4

/- ---+ / +,L-x

h--A-.-.-d

I

01 2

02 e

i 01% 0.2

X

+

Ae -0 %?!2 Wept

A e I Unswept Z

AQ Swept endualls -

Swept x Unwept +

oc 12”

X

/

+ +- --+-RR

x--x-, I ’

w(

-’ max

1

Fig.23 Variation of mass flow and pressure recovery with bleed exit area.

Intake on fuselage. LE vertical. &=- IO0

0.65

O-60

LE vertlcal wept endwalls LE vertwl unwept enduralls ----- LL horizontal wept endulalls -----

*a 6 e t max

*oo

0 A ef max

i

Pf 0 ii max

Pf

0 L C

Fig.24 Comparison of mass flow and pressure recovery. Intake on fuselage with LE horizontal and

vertical. 6, = --IO” (2) =O*IO

t

- 0.5

- 0.4

- 03

- o-2

- 0-I

0

- 0.4

- o-3

- 0.2

- 0-I

0

0 4 OL” 8 I2 0 4 ado 8 12

I I I

0 4 oc” 8 I

Intake on fuselaqe L I vertical

Wept endwalls 6, - IO”

-- 4 d” 0 I

Intake on fuselage L E vertical

tInwet enhualls 2-

e 0 0

Y Y Y Y

+

+ + +

X X X 1

1 1

4 oc 8 I2 A

0 $ O*lOi

Y Y Y Y :

X

+ : r

I I

4 ac” 8 I2

Intake alone 6, 0" Intake on fuselaqe (L E horizontal)

Fig.25 Engine face distortion coefficient .

ARC CP No. 1346 January 1973

533.697.24 : 532542.1 . 533.6 011.35

Brown, C. S. Goldsmith, E. L.

MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9

A rectangular mtake having variable geometry compression surfaces and destgned for super- sonic operation has been tested at a Mach number of 0 9 and a Reynolds number based on mtake entry height of approximately 10‘. Tests have been made on the mtake in isolation and on the side of a fuselage.

Maximum mass flow is less than e&mates based on geometrrc throat sire. The deficit corresponds to an effective reductton m throat area of about 3 per cent Cnticai pomt pressure recovery is lower than predicted, but the dtfference can be related to a ‘turnmg loss’ factor as was the case at supersonic speeds.

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ARC CP No. 1346 January 1973

533.697 24 . 532.542.1 : 533.6.011.35

Brown, C. S. Goldsmtth, E. L.

MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9

A rectangular intake havmg vartable geometry compression surfaces and designed for super- sonic operation has been tested at a Mach number of 0.9 and a Reynolds number based on mtake entry height of approximately 10’. Tests have been made on the intake m isolation and on the side of a fuselage.

Maxlmum mass flow is less than estimates based on geometric throat sire. The deficit corresponds to an effective reduction in throat area of about 3 per cent. Cnticai pomt pressure recovery IS lower than predrcted, but the difference can be related to a ‘tummg loss’ factor as was the case at supersonic speeds.

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ARC CP No. 1346 January 1973

533.691.24 : 532.542 1. 533.6 011.35

Brown, C S Goldsmith, E L.

MEASUREMENT OF THE INTERNAL PERFORMANCE OF A RECTANGULAR AIR-INTAKE AT A MACH NUMBER OF 0.9

A rectangular mtake havmg variable geometry compression surfaces and designed for super- some operatton has been tested at a Mach number of 0.9 and a Reynolds number based on intake entry height of approximately 106. Tests have been made on the intake in isolation and on the side of a fuselage

Maxmmm mass flow is less than esttmates based on geometrtc throat sire. The deficit corresponds to an effective reduction in throat area of about 3 per cent Crittcal point pressure recovery 1s lower than predtcted, but the difference can be related to a ‘turnmg loss’ factor as was the case at supersonic speeds.

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With the compressron angle of incidence wrthm however, there was some flow conditrons at angles removal of the swept endwalls.

The presence of thus partrcular the mtake.

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With the compressron surfaces of the mtake horizontal, performance was unaffected by angle of mcidence within the range 0” to 10” With the compressron surfaces vertical however, there was some loss in maximum mass flow and pressure recovery at crrtical flow condrtions at angles of incidence above about 4O removal of the swept endwalls.

This was only partly alleviated by

I With the compression surfaces t angle of mcrdence within

t however, there was some flow conditions at angles removal of the swept endwalls.

The presence of this particular fuselage imposed no obvrous effect on the performance of the intake. I The presence of this partrcular

the intake.

C.P. No. I346

Q Crown cepyright

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C.P. No. 1346 ISBN 611 470981 5