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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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