Oj

TRANSLATIONS FOR SCIENCE AND TECHNOLOGY

No. 324

Lingua Scientiae

Box 21086 Campus Station

Cincinnati, Ohio 45221

\ TESTS ON THE AFT INLETkMODEL,. SCALE 1:5, IN THE STANDARD WIND

TUNNEL OF THE DFL BRAUNSCHWEIG FROM 1 APRIL 1963 TO 28 APRIL 1963/

K. Knauer and K. Retti'g

7

? if

JAN "1

1- D

Translation of the report of Ernst Heinkel Fiugzeugbau GmbH,

Munich, dated 20 June, 1963.

January, 1966

i■: i./ii-OR^Aii'JW CENTER j

Translator's note

Due to the nature of this report, the translation will refer to

the equations in the original text without reproducing them. Marginal

annotation of the type /x will indicate the beginning of page x in

the original text.

The original version of this report contains

a) Text part

1 Title page

26 text pages

28 figures

3 table pages

total pages? 58

b) Supplement

1 title page

6 table pages

83 graph sheets

total pages; 90

page

1.

2o

3.

4.

5.

6.

7.

8.

9. •

10.

Summary

Notation

Evaluation of formulae

Inlet arrangement

Model and instrumentation

Test results

Index of graph sheets

Figures

Bibliography

Graph sheets

original

2

2

4

13

15

20

26

29

57

Supplement

trans

1

/

2

(p

7jO

/3

1. SUMMARY

A 1:5 scale model of the aft-inlet He-211 was investigated in the

standard wind tunnel of the DFL Braunschweig at flow velocities of

0-60 m/s, angles of attack of oc = 0° to oC = 15° and yaw angles

of (3 = 0° to /3 = + 15°, The results show very good agreement with

the theoretically obtained values. From this, it is possible to

attain a thrust gain of 6% at a flight Mach number of M = 0«8 and

25,000 ft altitude with the stipulated inlet arrangement as compared to

a simple Pitot inlet.

2. NOTATION

P (kg/m2)(kg/m2)

(m/s)(kg/m2)(kg s2/m4)

v

q

c

§ p (kg/m2)

m (kg/s)

*? ~ P /pS (kp)

W (kp)

M

FB = 15230 cm2

FT = 43.16 cm2

FF = 95.6 cm2

FR = 130.5 cm2

RB = 20 cm

RK = 18 cm

static pressure

total pressure

velocity

impact pressure

density

kinematic viscosity

pressure difference relative to the tunnel

static pressure

weight flow

inlet efficiency

thrust

drag

Mach number

reference area

area of engine test cross-section

area of fan test cross-section

area of exhaust pipe cross-section

total radius = fuselage radius

fuselage radius at the station of the

boundary layer test plane

/3

- 1 -

R0

<Pn

i

N

(cm)

(cm)

R (i - 10) radii at which the single boundary layer

probes lie

radius of the inlet flow tube at the station

of the boundary layer data plane

boundary layer thickness

(see figure 21 )

number of measuring points in the circum

ferential direction

number of measuring points in the radial

direction

total number of measuring points on the

circumference (engine, N=24; fan,

N=8; boundary layer, N=13)

I total number of measuring points on a

radius (engine, 1=6; fan, 1=4;

boundary layer, 1=10)

indices

t*? free stream

T engine

F fan

G boundary layer

R cross-section in exhaust

For clarification of remaining quantities, see "30 Evaluation of

formulae"

30 EVALUATION OF FORMULAE /4

In the following, the calculation procedure for the particular

quantities are given in the order that they appear from the result sheetso

Those definitions used in the data sheets are given in parenthesis,,

3«1 Tunnel impact pressure (Q-TUNNEL)

.equation (1)

302 Reynolds number (RE)

equation (2)

303 Density (RO)

3.4 Ratio of flow through the fan and engine (FLOW-RATIO)

equation (3 )

equation (4)

equation (5)

likewise

equation (6) '

equation (7 ) /5

— o —

3.5 Unsteadiness parameter Dc (DC60)

equation (8)

equation (9)

equation (10)

equation (11 )

where, besides, one is to set

n = 25 *- n = 1

n = 26 *» n = 2

n = 27 v» n = 3

D is, now,, the smallest of the 24 computed mean vlues DT

c60

equation (12)

After D , the <p-values for nQ and nQ+ 3 of the concerned Dy . are

stated on the data sheets.

306 Mean impact pressure (Q-MEAN) /6

qT from eqno (5); qp from eqn0 (6)

307 Mean impact pressure from orifice measurements! measured

ingine (Q-Blende 1), non-measured engine (Q-BLENDE 2)

equation (13)

equation (14)

3.8 Velocity ratio (G-Ratio)

equation (15)

equation (16)

3.9 Loss factors (LAMDA)

equation (17)

equation (18)

S p^ from eqn« (8); similarly S p0

3olO Mean standard deviation (OMEGA)

equation (19) •

similarly ' ' /Zequation (20)

D (n,i) from eqn0 (7); similarly Dp(n,i)

3.11 Unsteadiness parameter U (U)

equation (21 )

equation (22)

- 3 -

3.12 Total pressure coefficient

D..(n,m) from eqn. (7); similarly Dp(n5i)

3.13 Velocity coefficients

equation (23)

equation (24)

3.14 Mean impact pressure of the boundary layer flow through the

engine (QGM)

equation (25)

equation (26) /8

Under the assumption that the velocity profile is of the type

equation (27)

nn f

the quatities £ and n will now be determined from the above-calculated

values for each boundary layer rake by means of a quadratic error

adjustment.

From the continuity equation

equation (28)

one obtains, by consideration of eqno (27) and the substitution

equation (29) ,

the following equation for K:

equation (30)

For the mean impact pressure of a rake

equation (31)

one obtains, after the corresponding transformation for K£l: /9

equation (32)

for K>1:

equation (33)

These calculations can now be carried out for each of the 13 boundary

layer rakeso The ultimate mean impact pressure cfT is then given by

equation (34)

3.15 Boundary layer velocity ratio (GVG)

equation (35)

So 16 Boundary layer loss factor (LAMBDAQ)

- 4 -

equaLion (36)

equation (37) /lQ

equation (3S)

3,17 Boundary layer loss factor based on fan impact pressure

(LAMBDAF)

3,18 Mean distance of the stagnation point streamline from

the fuselage (IO-k-HOQ/R')For each particular boundary layer rake, R as well as h = RQ- FK f

can be ascertained from eqns. (29) and (30). / R.,

equation (39)

3-19 Fuselage drag compnent of the fan internal flow (CWF)

equation (40)

3o20 Loss factor of the individual boundary layer rakes (LAMBDA)

A~(n) from eqn. (37)

3,21 Distance of the stagnation point streamline from the

fuselage (10*HO/R)From eqns. (29) and (30), one obtains

equation

3022 Boundary layer thickness (DELTA)

3023 Boundary layer parameter (N)

See point 14.

3.24 Boundary parameter (H)

H = 2/N + 1

3025 cBoundary layer velocity ratio

VG(n5i)/Voo from eqn. (26)

3026 Determination of the inlet efficiency

For the determination of the inlet efficiency, it is necessary to

have a knowledge of the engine — particularly the fan airflow (my mp)

which is demanded of the engine, as well as the dependence of

these values on the pressure recovery,, In the present case, these

quantities were taken from the engine manual "General Electric CF-700-1

Turbofan Engine", April, I960- til

The calculation itself can only be performed iterativelys

1, Estimation of r> and np

-■ 5 -

Here, mk. signifies the flow at the estimated efficiency; m^O theflow at ^ , = o T = l.OO; i<T and Kp the correction factors given by

the engine.' manufacturer.

3. a* . /I!

M*

from'which

P / p = f(M*)i » o

V^/ V

4. From V^ / V and the test values, > is determined.

5. q

With the *7 -value (to be calculated each time) the calculation can

be repeated often, until a sufficient accuracy is attained.

Since both the engine and fan inlet losses influence the airflow

in the two regions, the influences of the engine inlet losses on the fan

flow and the fan inlet losses on the engine flow must be considered by

means of a higher order iteration.

4o INLET ARRANGEMENT /l3

The two inlets were designed especially for high-speed flight.

Auxiliary inlets were provided for take-off and landing.

a)Intake area: ¥,j = 0.105 m^ per engine. Figure 1 shows, for M = 0«8

and various altitudes, the sum of the thrust l@&& by total pressure

losses in the inlet ( &, S) and inl'et drag (W) as a function of inlet

size* Drag and losses were taken from reports [l][ and f.23 , in which

a systematic investigation of the Pitot-inlet is contained,, It shows

that the curves exhibit a minimum at F^y = 0.105 m , independently of

the altitude. From £l"] , the lip 1C can be chosen as the most favorable

lip shape.

Figure 2 shows the pressure recovery estimated from [l\ and [2J

over the range of Mach numbers. At rest and in low speed flight, under

the requirement that only 2,5% total pressure loss be permitted at MQ= 0,

the effective inlet opening must be increased to 0.185 m^ per engine bymeans of flapso '

Intake area: F = 0.253 m? per engine,,

- 6 -

lhe aft-fan inlet should, on the one hand, retard the boundary layei

entering the intake as .little as possible, on the other hand, permit a

not-overly-large inlet loss to appear. An inlet size was selected such

ihvi i the boundary layer accelerated up to a flight Mach number between

C.b and O.b, depending on the flight altitude, and was lightly retarded

at larger flight velocities (fig. 3). These inlet sizes gave about 2%

inlet total pressure loss over the entire high speed flight regime (fig. 4).

The calculation was based on an incompressible turbulent boundary

layer for a half-body corresponding to the aircraft fuselage (fig. 5),

so that the velocity profile /l4U/U1?O = (y/<S )1/7 . [3]

As with the engine intake, the inlet losses were taken from reports [l \

and [2j , whereby also the lip 1C proves to be the best here,, For 2.b%

pressure loss at rest, the effective fan inlet area must be enlarged to

Co350 m^ per engine*,

c ) c[rag_£_£ ._____ ^i_____

Due to the sucking of the fuselage boundary layer into the aft-fan

inlet, a body drag decrease appears on the one hand, while on the other

hand the net thrust due to the total pressure loss in the boundary layer

diminishes. Figure 6 now shows the effective drag decrease and net thrust

diminution. It indicates

/^Wp = m.(V^- V ) the drag decrease due to boundary layer(j V\j O , .

suction;

Sp net thrust of the engine with a Pitot

inlet having the same inlet loss as the

engine inlet;

S. Net thrust of the engine with aft-fan

inlet, with total pressure loss in the

boundary layer taken into consideration.

Moreover,

V free-stream velocity

V mean velocity of the stream tube reaching

°G from the inlet to the aft-fan entrance,with consideration of the acceleration or

retardation of the air through the inlet.

Airflows and thrusts were taken from the engine manual "General Electric

CF 700-1 Turbofan Engine", April, 1960,

5. MODEL AND INSTRUMENTATION /l5

5 o 1 MocI el

5,1.1 General

The problem existed, in the engine inlet investigations 6f the

aircraft He 211, to establish a truly similar flow in the' region of the

rear inlet and inside the inlet passage. The boundary layer on the

aircraft fuselage forward of the ring inlet, as well as the velocity and

pressure distributions in the engine compressor intake and at the aft-fan

- 7 -

section were Lo be invostiga tori „ In the inlet region, various

boundary layer thicknesses were to be produced.

From the construction plans, a 1:5 scale fuselage model was built,

with wing stubs and complete inlets, engine models and corresponding

i nstrumen ta t i on.

5.1.2 Model construction

The cylindrical central section of the fuselage of the aircraft made

segmented construction, composed of four cylindrical cast aluminum sections,

possible for the model. The forward section, the aft section with

empennage and inlet passages were of composite construction, made from

balsa wood and glass fiber-reinforced plastic. In order to produce various

boundary layer thicknesses in the inlet region, one or more of the

fuselage central sections can be removed, i.e. the fuselage length can

be shortened; the fuselage nose thus is drawn correspondingly toward the

aft (fig. 12). Figure 8 shows the model in the wind tunnel, as fig. 9

does also, but with shortened fuselage.

Especial care was devoted to the construction of the ring inlet and

the inlets in the empennage roots. The plastic construction made possible

the production of extremely smooth and true-to-plan inlet passages (figs.

11 and 13). The area variation of the rear inlet and engine inlet is /l6shown in figs. 14 and 15 (model measures!). For the determination of

the optimal fan inlet cross-section, its size can be varied by means of

interchangeable fuselage sleeve elements. Figure 16 shows the performance

of the inlet sleeves, which lie under the fan lip. The contour for the

triangular inlets in the fin roots is shown in fig. 17. Besides the

original fan lip cross-section, a second lip contour, which can be over

laid as a supplementary plastic piece, was investigated (fig. 17).

For the investigation of the inlet operation at zero freestream

velocity (engine operation in stationary aircraft) and small freestream

velocities (takeoff, roll), the inlet cross-section in the tail root was

enlarged, in order to obtain here a better airflow for the engine. This

was. done by opening lateral slits in the walls of the triangular canal.

These additional inlet area amounted to about 80% of the original opening

(fig. 18). For a part of the tests, an air deflector (as in fig. 18) was

built into the inlet channel, further to improve the pressure distribution

at the engine data section. From symmetry considerations, it was

permissible to measure the flow conditions on only one fuselage half or

in one inlet channel. Thus, one finds a half-crown shaped arrangement of

boundary layer measurement rakes (figs. 12, 18) on the left side of the

fuselage 0.33D (D = body diameter) ahead of the rear inlet, and in the

right engine duct the instrumentation for the determination of the flow

rate in the engine data section and fan data section (figs. 12, 20).

Details are in section 5.1.3. For the realization of a truly accurate

inlet flow, the engine flow in the two engines was simulated similarly by

suction. For each engine, the connecting pipes for the suction lines

(for separated engine suction and fan suction) to the pump are found at

the aft of the fuselage (fig. 10).

5.1.3 Model suspension ' /17The model is mounted in the wind tunnel by a three-point suspension.

The steel supports, which project out from the wing roots and 'connect to

the fuselage, were hung on the two hanger struts projecting down from

the tunnel balance platform, while the model rear was hung by a steel

wire from the balance (fig. 8 - 10).

lj,] .4 Instrumentation

The previously-mentioned boundary layer instrumentation consisted of

thirteen rakes distributed at equal angular intervals around the left side

of the fuselaqe ahead of the rear inlet. Each rake carried ten pitot

tubes, which were arranged more densely near the body than farther out

(fig. 11, 12, 18*).

In the engine data section, in the compressor intake plane, one

finds two total pressure rakes and two static pressure rakes, as well as

static pressure taps in the walls. Figures 20 and 21 show the construc

tion of the'model engine and the pressure rake arrangements„ The total

pressure probes are so spaced that each probe was assigned an equal-area

circular ring segment, which indicates a simplification of later

calculation operations*, In order to be able to obtain the pressure

distributions in the data section sufficiently accurately, the engine

hub was provided with rotating rakes. With the aid of an electric motor

built into the centerbody, the probe mounts can rotate stepwlse, and

thus the data plane can be scanned continuously with the probes* An

associated electronic positioning device furnished the remote control

and the adjustment of each selected angular position from the control

panel* The fan cross-section, at the station of the aft-fan compressor,

is provided with eight equi-angularly distributed total pressure rakes, as

well as four probes and eight static pressure taps in the outer walls.

All pressure measuring devices were connected with a multi-manometer /l8

in the measuring room by means of pressure hoses which were led out of

the model through the wing roots or the aft end of the fuselage.

5*2 Wiind_t unreel

The tests we.ro conducted in the normal wind tunnel of the DFL

Braunschweig (tunnel data: test section area 2.8 x 3.5 m, maximum speed

65 m/s, test section length 7 m).The balance platform located over the open test section was suitable

for the two-strut suspension of the model* The third suspension point

on the model rear was used to control the angle of attack by means of

raising of lowering a steel wire. Variation of the sideslip angle was

achieved by turning the entire balance platform. Two stationary suction

pumps, which were connected to the model by pipelines or flexible hoses,

served for'the production of the inlet flows. The maximum attainable

flow velocity in the data sections amounted to about 50 m/s. Fan or

engine flows were independently measurable and regulable,, The

quantitative control in the suction pipes was accomplished with the aid

of a remotely-controlled throttle valve* The angles of attack and yaw

of the model were controlled from the control panel.

A schematic diagram of the entire test arrangement is shown in

fig. 22.

5.3 I_n_s trumen_t£ti^ojn

The pressures taken from the data stations on the model were

transmitted by hoses to two multi-manometers, so that the pressures of

fan and engine data planes were displayed on manometer A, the boundary /l9

layer pressures on manometer B. The manometer contains a contrasting-

colored liquid alcohol ( )f - 0.184 kg/dm ) and were photographed (2

* translator's note; This should read fig- 19, rather than 18

- 9 -

robot cameras with 10 m pi at olio 1 ders) for the analysis of the test values..*

Figures 23 and 24 show manometer photographs, as they were analyzed in

negative slide format. As was mentioned in section 5,2, the amount of

air for engine flow simulation was measurable separately for each line*

The measuring occurred with the aid of Pi tot tubes and static pressure

wall taps in the suction lines. The measured pressure differences were

displayed simultaneously on the control manometers on the control panel

and on the multi-manometers, just as the tunnel ram pressure-

For the remote control, rotating test probes in the engine data

section show that test positions in the radial section at 15° intervals

to be sufficient; this test arrangement gave a picture series of twelve

photographs per test condition.

5.4

The film negatives (24 x 36) with the test data photographs were

read with a semi-automatic analyzing machine (Telerecorder, Fa. Data

Instruments), the test data put on punched cards, which were further

processed on an IBM 7070 data processing machine. The form of data output

is shown on the enclosed tabular sheets (figs. 25, 25a), The explanation

of the symbols appearing therein is to be found in section 3.

6. TEST RESULTS . /20

6.1 in.£.liiel!c£ of the boundary layer

The effect of the various fuselage lengths on the characteristic

values of the inlet is represented In sheets 4 through 12 of the

Supplement, Theoretical investigations show that the potential flow in

the environs of the inlet is practically independent of the fuselage

length. This can be traced back to the fact that the body is foreshortened

merely in its cylindrical part. One can therefore consider the consequences

of the flow distance variation-induced changes in the boundary layer

thickness as the exclusive cause of the experimental differences in the

inlet inflow.

The loss factor for the engine inlet (sheet 4) does not, as expected9

remain constant, but increases with increasing boundary layer thickness,

especially at higher speed rates. This can be explained as follows; At

higher rates of speed, at which the engine and fan inlets are very

strongly choked, a more or less greater part of the boundary layer is

deflected from the body, reaches the engine inlet and develops there a

strong deflection jointly with the initiation of separation. This is

all the more stronger for thicker boundary layers and more completely

choked fan inlets, A further proof of the flow separation is to be seen

in the increase of the unsteadiness parameter (sheets 7, 9 and ll)o

As was to be expected, the fan losses (sheet 5) increased with

fuselage length, for with increasing boundary layer thickness more and

more air attained lower energy in the inlet. The pure inlet losses

(sheet 6) hardly varied thereby,

* translator's note: It seems unlikely that these are the correct units.

It is conceivable that 10 "m-Kasette" means 10-exposure film cartridges.

- 10 - v

The fuselage darg, which is contained in the internal flow (sheet

12), is about 20% higher than that given by theory«» The reason is to

be seen in the strongly turbulent flow behind the struts, which was /21not taken into account by theoretical means. This indicates that,

through the existing inlet arrangements, a part of the wing-body inter

ference drag is also compensated.

6.2

The effect of angle of attack is shown in figs. 13 -through 21 of the

Supplement, the effect of yaw angle in figs. 22 through 29. Up to <*- = 6°,

no variation is established with respect to the total pressure loss and

the flow unsteadiness in the engine inlet. At yet greater angles of attack,

the loss rises sharply. The turbulence, on the other hand, only becomes

markedly worse at oc = 15°. This indicates that at cruise condition

(Vco /V-r- ^ 2), a marked thrust decrease is to be expected from the engineinlet at c< > 6°, and beyond <* > 10° an angle of attack limitation is

to be imposed because of excessive flow turbulence.

It is otherwise with the fan inlet. Whereas here the total pressure

loss decreases with angle of attack, the unsteadiness due to the unsymmetric

flow already increases markedly at oc = 3°o Since the exit level ( oc = 0)

is significantly lower than the engine inlet, the turbulence is smaller

over the entire angle of attack range than for the engine inlet at oC = 0.

Yawing the flow on the fuselage to the inlet side ( fi = - 5°) gave

an improvement of the pressure recovery and the flow turbulence, at

(3 = + 5° a deterioration. With two built-in engines, these effects

will nearly compensate one another. Thus, in the investigated yaw angle

region, the thrust decrease due to yaw was not calculated.

The exact locations of the limits on angles of attack and yaw /22imposed by the unsteadiness can only be given after the completion of

a comprehensive engine calculation.

6.3

Sheets'30 through 34 of the Supplement show the various inlet values

for the two lips tested. At rest, both the fan inlet loss and the

turbulence parameter were reduced up to about 30% by means of the thicker

lip (lip 2). From V^ / Vp = 1.0, the two lips can be seen to be equivalentin relation to the internal flow.

6.4 E.ff.ect^qf £uj>ejLa2e_C£nj3triiCti>on>

The fuselage constriction (figs,, 35 through 43 in the Supplement)

did not bring the really expected effect. The diminution of the fan

inlet loss at rest for the most severe constriction was, indeed, of the

same order of magnitude as that obtained with the thicker fan lip, but

the turbulence was not essentially improved. The reason for this is

presumably to be sought in the constriction of the flow shortly behind

the inlet lip, and an associated separation«

The constriction had an unforeseen effect on the engine inlet.

With constriction 2, the total pressure loss in the engine was substantially

reduced in the higher velocity range. It increased again with more severe

constriction. The unsteadiness remained nearly invariant. This can only

be explained in that, with larger constriction, over-velocity appears on

the interior lip of the engine inlet, which leads to losses in the subsonic

diffuser. The values for fuselage length 3 (Supplement sheet 64) show

- 11 -

that this effect appears only for the Mucker boundary layer (fuselage 1). /23For the thinner boundary layer (body 3), no influence of the constriction

on the engine inlet exists,

6,5 AuxiJJ-a/y ^nr^ijno"^Q^i2^

VV}~ii 1 o the supplementary engine inlet (fig. 18) should theoretically

bring a reduction in the stationary losses of up to about 70%, an

improvement of merely about 17% can be ascertained from the tests.

(Sheets 44 to 47 in the Supplement) The reason lies in the overly

small cross-sectional area which the subsonic diffuser exhibits to the

auxiliary inlet at altitude. Also, a deflector plate built into the

subsonic diffuser brought no further improvement (Sheets 48 to 51 in the

Supplement).

6 . 6 S_imi_]_a£ii1y_pa_ra_m£t£r_s

Reynolds number independence was assumed for the conversion of the

test values to the large model. This means that all quantities referred

to the mean ram pressure in the engine or fan are directly valid for the

large model. The essential similarity parameter is the boundary layer

thickness referred to the body diameter. Since the measured boundary

layer varied strongly around the circumference, due to deviation of the

fuselage from rotational symmetry and the distortion of the flow in the

wing wake, the boundary layer thickness at the fan inlet was determined

theoretically both for the full-scale version and the model at various

body lengths. In fig, 7, one has plotted the ratio of model body lenth to

full-scale Reynolds number for which the ratios S /Rn for model and

full-scale version are equal. Thus, the one body length corresponding

to a full-scale Reynolds number can be assigned.

Sheets 75 through 85 in the Supplement show the dependence of 7* y,

>\ p and Cmjp on Mach number and altitude. On the left half of the sheets, /24the dependence of the corresponding inlet values on the Reynolds number;

the dependence of Reynolds number on Mach number and altitude is given

on the righf side. Thus, with the given Mach number and altitude, the

appropriate inlet parameter can be read off directly.

6,7

In fig. 26, the pressure recovery is plotted for the Mach number

dependence of the General Electric CF 700 engine at various altitudes.

The calculation was based on the inlet quantities measured with con

striction 1, fan lip 1 and closed engine inlet. Comparison with the

predicted values (figs, 2 and 4) shows that the engine inlet is somewhat

poorer than expected in the lower Mach number range,, If the high

stationary losses cannot be accepted in the bargain, either the inlet

must be enlarged or the inlet lip thickened. An auxiliary inlet appears

to be unprofitable, due to the only slight improvement attained therewith*

The fan inlet loss is significantly less than expected in the lower

Mach number range* A further improvement may be attainable at all Mach

numbers with fan lip 2 and fuselage constriction 2,

The thrust increase attained compared to a Pitot inlet is somewhat

larger than expected (fig. 27), At a cruise altitude of 7 - 8 km and

a Mach number of 0,7, it amounted to about 8%.

All in all, the tests have confirmed the advantages of the He 211

inlet compared to the conventional inlet arrangement. Besides the

- 12 -

considerable thrust increase, there is, moreover, to be expected an

improvement of the base aerodynamics of the aircraft, not investigated

here. The remaining difficulties (strong flow turbulence and high

stationary losses in the engine inlet) can be overcome easily by

appropriate variations*

7. INDEX OF GRAPH SHEETS

(Column headings)

Column

1

5

6

7

/2k

Legend

Sheet number

Fuselage length number

Fan lip number

Constriction number

(footnotes)

* without deflector, inlet slits open/closed

*■* without/with deflector, inlet slits open/closed

/27

8o FIGURES (legends only) /29

1,

2.

3.

4.

7,

10.

11.

12.

13.

14.

Sum of thrust loss due to inlet loss and drag, dependence on

inlet size.

He 211 engine inlet. Assumed pressure recovery*

(dashed line) without auxiliary opening

He 211 aft-fan inlet. Distance of the stagnation point streamline

from the bodyc (data symbols for four different altitudes)

He 211 aft-fan inlet* Assumed pressure recovery,,

(dashed line) without auxiliary inlet; Grenzschicht = boundary

layer; Einlauf = inlet.

Impulse defect thickness and boundary layer thickness on a

rotationally-syrnmetric half-body with turbulent flow at the

station x/R = 15,9 (incompressible)He 211 aft-fan inlet. Effective drag reduction due to suction

of the fuselage boundary layer,, (data symbols for four altitudes)

Station at which the impulse defect thickness on the 1:5 scale

model is equal to that of the full-scale model at x/R = 15.9(Calculation for rotationally-symmetric half-body, incompressible)

Model flow velocity = 60 m/s; model data points for various lengths.

He-211 rear inlet model in the wind tunnel.

He 211 rear inlet model with shortened fuselage,,

He 211 rear inlet model, view of the test arrangement*

He 211 rear inlet with boundary layer rakes«

He 211 detail. Inlet area per engine? main.: inlet — 0«105 m2,aft-fan -- 0.253 m

2 (Rumpf = body, schnitt = cut)

Main inlet (most of the legends are so reduced as to be illegible)

He 211 area variation (inlet rear) of the aft-fan

upper curve: area without engine inlet channel

lower curve; true area

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___ -V

15. Ho 21.1 engine inlol .»nvi variation, normal to the flow,

16. Ha]f~ring for i ■ 111 ..kIjik, I men I.

C ou n L o r s ij n k hoi e r, a t 1 5 ° off s c t.

FUncj fit in soclion piano and surface area adapts to. body

surf a ce s t r e a m 1 i. n e .

EinschnUrung " constriction; ringflache ~ ring area

17. He 211 inlet lips (MACA RM L56C28)

Lippe ~ lip; einlauf ~ inlet; schnitt ~ section; achse = axis,

18. Arrangement of the auxiliary slits and the inlet baffles in the

rear inlet of the He 211.

19. Body data region, He 211.

• Spant = bulkhead; Rurnpf = body; Blatt = sheet; Messrechen = rakes;

Triebwerksachse = engine axis; nach Montage gemeinsam abfrasen =

jointly milled after mounting; Aufnahmering = mounting ring.

20. Complete right engine (He 211).

geklebt = glued; Ansicht in Richtung Y = View in direction Y*

21. Test probe distribution, He-211.

Grenzschichtrechen- boundary layer rakes; Druck - pressure;

statischer = static; gesamt = total; Triebwerk - und Fanrechen -

engine and fan rakes; Ansicht von vorne gesehen = view seen from

the fronto

22* Test arrangement in the wind tunnel,

Sondenstellungsgeber = probe positioning device; Bedienungstisch =

control platform; Drosselklappen = throttle valves; Kanalsteurpult

= tunnel control panel; Durchsatzmessstellen = flow measuring

station; Druckmessleitungen - test pressure lines; Absaugleitungen

= suction lines; DiUse ■- nozzle; Auffangtrichter = collector,

23. Multi-manometer with engine and fan pressure display,

24. Multi-manometer with boundary layer pressure display,

25a. Evaluation of subsonic fan inlet tests (boundary layer)

(lower table): Boundary layer flow ratio

25. Evaluation of subsonic fan inlet tests,

Gesamtdruck-beiwerte = total pressure coefficient; Triebwerk =

engine; Geschwindigkeitzahlen = velocity numbers,

26. He 211 rear inlet model, 1:5 scale, pressure recovery as a function

of Mach number for various altitudes,

Vollastschub = Full-load thrust (max thrust); maximaler Dauerschub

= maximum continuous thrust,

27o He 211 rear inlet model, Is5 scale, effective thrust increase

of the present inlet arrangement compared to a Pitot inlet,

SUPPLEMENT

Blatt = sheet

1. Rear inlet model He 211 — graph arrangement*,

■ Versuchsnumir.er = test number; Rumpf = body; Fanlippe = fan lipj

Einschntlrung = constriction,,

2, + 3. same as for L

4O Data section •: engine,,

Rumpflange = fuselage length

5. Data section; fan

6o Data sections boundary layer and fano

7. Data sections engine.

8o Data sections fan.

Legends for the following sheets are identical with 6°. 15, 24, 31, 379

54, 61,

/£!

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Legends for the following sheets are identical with 7: 9, 11, 13, 16,

18," 20, 22,25, 27, 29, 35, 38, 40, 42, 52, 55, 57, 59, 62, 64,

67, 69, 71, 74.

Legends for the following sheets are identical with 8s 10, 12, 14, 17,

19, 21, 23, 26, 28, 30, 32, 33, 34, 36, 39, 41, 43, 53, 56, 58,

60, 63, 65, 66, 68, 70, 72, 73.

44o Data section: engine,

inlet slit open

inlet slit closed without baffle

45. Same as for 44

46. Data plane: engine,

inlet slit open, without baffle

inlet slit closed, without baffle

47« Same as for 46

48, Data section: engine,,

inlet slit open

inlet slit closed

. . inlet slit closed without baffle

49o Same as for 48

50. Data section: engine*

inlet slit open with baffle

inlet slit closed, with baffle

„ inlet slit closed, without baffle

51. Same as for 50

75o Rear inlet model He 211, '\j as a function of Mach number and

altitude at various velocity ratios.

76o Rear inlet model He 211, AF as a function of Mach number and

altitude at various velocity ratios.

77. through' 84-: Same as for 76

85. Rear inlet model He 211. Cwp as a function of Mach number and

altitude at various velocity ratios.

86o Same as for 85

BIBLIOGRAPHY " ' /57

1. Transonic Wind Tunnel Investigation of the Effects of Lip Bluntness

and Shape on the Drag and Pressure Recovery of a Normal-Shock

Nose Inlet in a Body of Revolution., NASA RM L56C28.

2. The Effects of Lip Shape on a Nose-Inlet Installation at Mach numbers

from 0 to 1.5 and a Method for Optimizing Engine-Inlet

Combinations, NACA RM A54B08.

3. Charts of Boundary-Layer Mass Flow and Momentum for Inlet Performance

Analysis. Mach number Range 0.2 to 5.0. NACA TN 3583.

4. Ehrismann, Aos (Unsteadiness parameter Dq for the Total Pressure

at the Compressor Intake of TL-engine5? (28 Nov 1962))5. Ehrismann, A,: (Symbols and Definitions of Aerodynamic Quantities

for Inlet Tests. (3 Oct 1962))

6. Soners, H. Do: Boundary Layer Ingestion. Design Information Memo

randum #401. Flight Propulsion Laboratory, General Electric

Company, Cincinnati, Ohio.

7. General Electric CF 700-1 Turbofan Engine. April, I960.

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88 Truckonbrodt, Eo; (A Quadrature procedure for the Calculation of

the Laminar and Turbulent Friction Layer in Plane and

Rotationally-symmetric Flow. Archive no, 209 pp* 211-228,

1952.)

VIV i

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