NASA
Contractor
Report
An
Experimental
the
Aerodynamics
3405
Investigation and
Horizontally-Opposed Aircraft
Engine
Stan
and
J.
Texas
Miley A&M
College John
Air-Cooled
Installation
J.
Cross,
Jr.
Texas
Owens
Mississippi
State
Mississippi
State,
David
L.
Turbo
West
University Mississippi
Lawrence
Broomfield,
Prepared
of a
University
Station, K.
Ernest
Cooling
of
Corporate
Aircraft
Center
Colorado
for
Langley Research Center under Grant NSG-1083
N/LSA Nationa_
Aeronautics
and Space Scientific Information
Administration and Technical Branch
1981
,1_
-¸
SUMMARY
A
flight
test
investigate
the
based
research
aerodynamics
opposed
aircraft
engine
tigated
were
internal
of
the
installation,
dynamics. art
The
are
the
cooling
to
established
a
flight
ments The drag
is
of
much
of
for
engine
terms of
shown
of
for
inlet
and
to
of
be
presented.
internal
radial
aerodynamics
engines
are
Correlation
the
installation.
particular
development
exit
design
test
installation
engines.
measurable
major
ground
cooling
the
easily
and
manufacturer's
of
method
aero-
state-of-the-
cooling
the
inves-
mechanics
exit
test
are
areas
cooling
current
the
to
a horizontally-
and
Flight
problems
that
measurements
impact
area.
horizontally-opposed
test
in
and
developed
between
flight
theory
cooling
technology
and
aerodynamics,
each
show
of
performed
Specific
aerodynamics
development
of
results
applicable
and
the
solution
The and
for
was
cooling
installation.
applicable
for
and
inlet
discussed
techniques and
the
program
of
is
design
cooling
data Also,
require-
parameters
is
presented.
on
and
cooling
cooling
significance.
INTRODUCTION
The
research
established cooling
drag
to
program,
perform
associated
an
which
is
exploratory
with
reported
herein,
investigation
reciprocating
engine
was of
powered
the
general
aviation
apparent
that
aircraft. attention
cooling
installation
cooling
drag.
engine
cooling,
altitude,
also
for
but
related design
also
inadequate
result
Cooling behavior cooled
fuel
high
The first
three
geometry
requires
engine
consisting
four
the
plenum,
components
perpendicular
geometry,
concerned
for
however, primarily
the
basic
in-line,
airflow
by air-
engine geometries
the vee,
and the
consisting plenum, because
of an
and exit. the
engine
to pass through path.
a relatively
of a cylindrical
the
system configurations.
are necessary
to the flight requires
cooling
engine
low pressure
cooling
with
The components which
use the same system,
pressure
at
both
on tbe particular
airflow
geometries,
horizontally-opposed,
is
engines.
dependent
cooling
The standard
the engine
system required
1 illustrates
cool-
efficiency.
aerodynamics
system are
four
in excessive
Consequently,
fuel
at
aerodynamics.
results
flows.
aircraft
and the associated
inlet,
engines
supercharged
to operate
reciprocating
Three of the
for
cooling
in reduced
Figure
inadequate
only
is
as
area of concern,
in poor cooling.
of the airflow
geometry.
not
became
are manifested
to installation
installation
make up the
which
it
on the engine
can result
higher-than-necessary situations
be focused
aerodynamics
particularly
Poor aerodynamic
cure
should
An associated
is
ing drag,
As work progressed,
The radial simpler
cowl.
the engine
system
The airflow
passing path.
through
the engine
The cowl
functions
airflow
through
captured corporates
the
Cooling areas.
cooling
flow
through
heat
transfer
radial
the
flow
technology
concerned
is
with
area is
concerned
only
design
of little
survey
was performed
cooling
flow/heat
internal
survey
will
this
time,
powerplants, tially
problem
be published
and mainly interest
developed
is
dominated
the external
radial
geometries.
engine
installations.
material
literature
applicable and cowl
Over five
The results
external hundred
All
literature
report.
by the development
engine
to the
of the
related aspects
of radial
1930-1945.
to the development
from print.
cowl/nacelle
an extensive
areas.
and
The existing
encompasses the period turned
for
to horizontal-
to similar
as a separate
and air-cooled
disappeared
with
aerodynamics
have been collected.
The literature engines
transfer
aerodynamics
references
of the
and vee geometries.
program,
to identify
in-
two problem
use to horizontally-opposed
of the research
the
and the resulting
system.
amount of data regarding
As part
involve
applicable
as in-line
applicable
also
the character
also
installation
forcing
It
The technology
area is
flight
structure.
fins
The large is
baffle,
its
as well
of the
here
in
cooling
flow.
in this
The second problem
exit
to the
fins.
aerodynamics
engine's
engines
internal
the coolin_ flow
to this
engines
ly-opposed
is
parallel
as an external
installation
The first
remains
After
of gas turbine
technology
essen-
of the problem
areas
are well
covered
the non-radial tions
_eometries.
during
erences
the radial
1-6)
These are,
concern
for
tions
cowl
also
One ground
the
was also the
The remaining
programs
Avco
the
Avco Piper, and
the
wish
for
and
to
the use
in
by
and
Teledyne
engine
of
is
the by
attempt-
of
this
Piper
Continental
for
and
Aircraft, appreciated.
representatives American,
con-
research
aircraft
greatly
the
aircraft.
important
to
test
aircraft.
aspects
the
program
Grumman
asset.
program
industry
donation this
test
engine
twin
engineering
Cessna,
important
test
acknowledge
Propeller
Beech,
an
a
program
on the
various
aviation
Hartzell
participation
was
installa-
of flight
a single
utilizin_
general
system
Rockwell,
review
years,
investigation
investigated
particular,
Lycoming,
to the
the research
The first
drag of
aerodynamics
Lycoming,
Also,
cooling
investigators
propulsion
test
conducted.
ed to measure
In
(ref-
7-11).
investigations.
program.
six
to horizontally-opposed
a series
by
cita-
aerodynamics.
In recent
and involved
tributions
of
plus
only
applicable
was experimental
The
hundred
or vee cowl
survey,
installation
aerodynamics
period,
literature
aircraft
for
five
geometry.
has been given
Except
the cowl
development
the most part,
(references
for
Of the
in-line
horizontally-opposed attention
except
from
Mooney, program
critique
SYMBOLS a
effective
a 1,2,3
b
coefficients equation (5) exponent
C 1,2,3,_
Cp
orifice
of
constants equations pressure
of
area, least
orifice
m 2 squares
power
used
in
used
in
law
of heat transfer (8) - (II) coefficient
surface
P
power
laws
- P= qco
H
heat
I
indicated
k,l,m,n
exponents of power coolin_ correlation (8) - (13)
P
static
P_
free
stream
static
CO
free
stream
dynamic
T
transferred engine
ambient
TEGT
exhaust
T
temperature engine, °C
T g
effective
(free gas
time,
joule/sec
kw
laws in development relation, equations
of
N/m 2 pressure,
stream)
Th
cylinder
Th A
average of the 6 temperatures, °C
Th 6
cylinder head number 6, °C
N/m 2
temperature,
cooling
combustion
N/m 2
pressure,
temperature,
of
bead
unit
power,
pressure,
a
ex
per
°C
flow
gas
temperature,
engine
temperature
°C
exiting
temperature, °C
cylinder
of
head
cylinder
the
°C
T up
temperature at rear of
T*
temperature
ratio
heat
transfer
temperature
ratio
of
of
Th6
W
for
charge engine,
C
cylinder
cylinder
air
mass
flow of kg/sec
flow, fuel
coordinates used determination of equation (5), m
ap
engine
baffle
density
head
plenum,
based
on
temperature
ratio
based
on
use
temperature
approaching
m/sec
and
air
through
the
for integration in the cooling air mass flow,
kg/m
ratio
inlet,
kg/sec
pressure
density,
upper
m/sec
x,y
air
head
temperature
velocity
cooling
W
for
velocity,
flow
O
ThA
transfer
inlet
V
in
transfer
heat
1
flow
heat
use
Vo
of cooling engine, °C
drop,
N/m 2
_
relative
to
standard
sea
level
density ratio of cooling air flow exiting the engine relative to standard sea level
°ex
MEASUREMENT
The reliable
objective measurements
using
flight
would
attempt
stallation
of
test to
OF
this of
program
cooling
techniques. identify
components
to
COOLING
the the
was drag
If
so,
DRAG
to could then
contributions cooling
drag.
ascertain be
obtained
follow-on of
if
testing
specific
in-
The aircraft (Figure
2) obtained
installation in that
exhaust
leading
ejector
through
propeller
feathered,
All
flights
to witbin
one knot.
meters
altitude
in
Configurations closed,
closed.
unique
of the
feathered
gliding
for
this in for
was required
The results
figure
near
i.e.,
full-feather-
morning
good data
runs were kept
of at
least
an acceptable
included
tubes
tubes
of the
stopped,
the early
were tested open,
were made
program.
glide for
the
engine
A special
A stabilized
ejector
in the
sinks,
practice
are used at
fuselage,
flight.
deviations
which
The cooling
by current
are visible
were performed
Airspeed
U.S. Navy.
The drag measurements
was installed
calm air.
Table
tubes
of
was a Beech T-34B
pumps, or augmentors,
the bottom
technique
ing propeller
lets
is
edge of the wing. the
program
from the
aircraft
The ejector
extending
this
on loan
of the
the exit.
usin_
used for
inlets
1,500 data
open,
restricted,
drag measurements
and in
run. in-
and tubes
are given
in
through
the
I. The drag associated
installation
is
aircraft
drag.
of the
internal
no-flow
drag.
11 for
a twin
with
indicated
the cooling
to be seven percent
Drag calculations flow
yielded
These values engine
flow
of the no-flow
based on the momentum loss
a value are similar
of six
percent
to those
of
the
of reference
configuration.
7
As a result
of this
became evident resolvable
that
through
investigate
test.
of
was abandoned
aerodynamics.
by good aerodynamic emphasis
design
of the research
trial-and-error
scientific
approach
various
aerodynamic
cooling
installation.
of
interest
in favor
of
program
the cooling
program
thus
studying
the
here is
that level
installation.
shifted
from
The a tradi-
drag clean-up
approach,
investigating
and understanding
effects
involved
to
installation
to the minimum necessary of
it
were barely
of specific
The point-of-view
drag can be reduced
program,
The follow-on
the drag contributions
associated
tional
drag measurement
drag levels
flight
system components
cooling
the
first
in the
to a more
operation
the of the
COOLING INSTALLATION AERODYNAMICSPROGRAM Test Aircraft The test prototype 3.
pressurized
The engines
of 201 kw. in excess for
test
aircraft
and Cooling
used in the program Aztec.
The aircraft
were turbosupercharged
The aircraft of
Installation
was capable
7,000 meters.
the various
studies
is
PA-41P
shown in Figure
a sea level
of operating
The starboard
rating
at altitudes
power plant
was used
of the program.
A schematic
of the cooling
aircraft
shown in Figure
is
with
was a Piper
installation 4.
Particular
of
the PA-41P points
of
reference
are denoted
by numbers.
of most horizontally-opposed
engine
of cooling
flow
been acted
on by the propeller
stream
is represented
conditions.
is
conveyed
the
cooling
exhausted
through
the cooling
the exits
air
flows
known as downdraft are also
the
cowl,
although
the
upper
surface.
is
engine
specified
requirements
Figure right in
5.
terms
temperature ordinate.
and the The
side of
to
of
exist
with
coolin_
requirements
are
temperature, the
engine
operation engine
required orifice
where
of
is
to top.
the
the
flow
of
exits
on
cooling to
The
cool
determined
characteristics
to
from
condition
is
normally
similar
and
is
read
refer
the
specified
cylinder
then
the
airflow
are
a form
power,
airflow
plenum,
surface
characteristics in
The
and is
and Design
air
manufacturer
graph.
through
from bottom
conditions.
orifice
and
configurations
lower
sufficient
(2)
(4)
on the
engine
(CHT); The
cooling
function
operating
engine
the
air
conduct
inlet
configuration,
Operation The
the
plenum
flows
generally
requirements.
under
air
Installation
installation
by
Updraft
has
at free
to the lower
some arrangements
Cooling Cooling
is
This
the upper
where the
of exits
The flow
then flows
lower
(5).
from
It
typical
The source
(I).
enters
(3).
the
cooling.
available,
The location
supplied
air
into
is
installations.
and is no longer
to the upper plenum passages
arrangement
by point
The cooling
fin
This
head
from to
the
the
relationship
between
passages
pressure
which
relates
decrease to the
most air-cooled Theory cooling
of
the
free
to
the
airspeed
pressure,
cooling
The
engine to
the
air
air
static
I0
control
the
the
the
is
flow.
and flow
external
rate.
to
head
If then
the exit
be
upper
plenum
part
of (3)
at
dynamic
the
the
the to
air
head. of is
exit
region
of
pressure
(5)
The
such
that
(6)
a hinged
flap, be
used
(4),
external
can
the
flow
accordingly.
local use
should
upper
auxiliary
passages
changes
the
The
and
the
the
static
fin
region
(i)
captures to
engine
of
air
pressure
plenum.
cooling
through
Through
the
the
(2)
point
the
corresponding the
inlet
the
of
cooling
head
recovery
from
equal
the
upper
this
density
accelerates
pressure
area
drop,
schematic
dynamic
the
at
(7),
passing
of
The
(4).
cooling
and
the to
plenum
heated
of
this.
used on
actuality,
for
full
through
lower
In
pressure exit
with
path.
then
In
to
air
a
pressure
a reservoir
proceeds
the
source
the
flow
cooling
auxiliary
is
fin
accompanies
plates
gives
propeller.
the
as
6
The
converts
conveys
The
then
this
heated its
partially
into
the cooling
pressure
baffle
aircraft.
the
conditions
supply
takes
by
flow.
flow
which
the baffle
a dynamic
the
serves
stagnation
called
Figure
with
of
and (3)
is
model.
stream
air,
in pressure
operation.
modified
coolin_
through
engines.
is
is
of flow
intercylinder
installation
plenum
rate
and the decrease
This
head
the
static
varied
both
In actuality,
the pressure
may be as low as fifty pressure,
percent
due to flow
losses
sufficient
plenum volume
sequently,
the
of the
results
is not
flow
through
the cooling
passages
Cooliq$
installation
design.
given
in
Also,
con-
over flow
engine
efficiently.
Procedures
for
12.
the
to vertical
accomplished
7 and
in-
velocities,
from horizontal
references
plenum
dynamic
inlet.
distributed
nor
are
the
stream
in finite
evenly
the transition
in the upper
free
through
face
analysis
is
flow
recovery
They
the are
design summarized
here. The
design
of
the
cooling
installation flow
dimensional
subsonic
compressible
the
pressure
head
dynamic
Figure
6],
applying
the
a pressure
recovery
factor
diffusion
developed
losses
incurred
whether
should
be
section. baffle flow
This
used The
pressure rate.
altitude is
the
As
between
static
here.
This
information
The the
or
flow
(4)
5,
is
in
been
and a
total in
(I)
a
pressure
across
lower the
supplied
in
by
for
the
the
flow as
pressure later
determined
the
with
question
required
the
rise not
is
the
the
density
has
(2)
addressed
with
Figure
accounts
There
one-
[point
determined
inlet
(3).
be
is
on
Starting
inlet
which
the
temperature
heretofore
(3)
pressure
in
the
pressure
associated
indicated
from
and
will
plenum
drop
dependent.
determined
(I)
plenum
lower
by
based
theory.
of
pressure
of
to
front
plenum
amount
upper
in
is
by
the
cooling drop
air is
plenum engine. by
engine
II
manufacturers,
and estimates
Typical
values
range
is
sized
then
pressure exit
is
the
and associated
the
same as the
flow
local
system throttle
pressure
drops will
the
expanding
so that
the
area
exit
that
adjust
pressure.
its
area
(6).
The
the flow
so that
the
For a given
increases
(5)
static
flow
in
external
must be used.
The exit
external
as the
matches
condition,
50°C to 70°C.
to accelerate
area acts
pressure
from
based on experience
rate exit
flight
the flow
rate,
and conversely. The design
problem
is made difficult
in horizontally-opposed prime With
concern the oil
tend
ter
of the flow
the
location
governor.
tion,
in
losses. for
which
requirements 12
the
the
cooling
induction
of the requirements configuration
between
create data. is
The characby
and exhaust
alternator,
of pressure
requirements
in-
is affected
to some engine
differences
configuration
test
on top.
of an inter-cooler, leads
engine,
plenum volumes
regions
of the
the determination
the installation application
respective
This
Also,
large
Of
plenum volumes.
of the
plenum volumes
and configuration
empiricism flow
in these
and the presence
propeller
on the bottom
small
variation
configurations.
and lower
to have relatively
and relatively
lines,
engine
here are the upper sump located
stallations below
aircraft
by the wide
and/or
dependent
recoveries
and
the test
configura-
are determined, uncertainties
and
as to the
A typical
cooling
shown in Figure
7.
This
is
the
ideal
exists
cooling
on top of the
uniformly
across
the cooling measure
air
is
temperature
of the
rear,
and the
that
of Figure Cooling
in the
operation
in
context
previous mechanics,
eral the
of the
pressure
is
the
highly
this
industry. orifice
is
of a true
installation nonuniform,
progresses
towards
the the
plenum pressure
and
Investigations Aerodynamics
aerodynamic
and exit with
measurements
design
effects.
problems
in the
internal
areas.
flow
The first
methods of instal-
in current
The second problem
engine
was done
The internal
of the various
and test
involved
This
two problem
characteristics
Program
effects
installation.
measurement
was drawn from the radial performed
For
Three areas were studied:
of an evaluation
installation
as it
aforediscussed
dealt
engine
flow
of the cooling
investigation
aviation
drop.
Installation
the various
effects,
flow
distributed
The temperature
and Aerodynamics
inlet
consisted
is
and the plenum pressure
between
of the
section.
face.
air
plenum
open to question.
Installation
was to investigate
engine
rises
relationship
a true
and the cooling
4, the
flow
7 is
in that
pressure
Figure
The objective
lation
uniform
of the baffle in
the
engine
the upper
configuration
flow
configuration
use by the
gen-
area dealt
with
and the correlation
between
cell
Much
cooling
measurements. correlation
work
by NACA.
13
The inlet some basic
and exit
design
Four different of both
cowl
parameters
inlet
flow
investigated
flap
investigated
the effects
on installation
configurations
the external
parameters
studies
included
performance.
were studied
and internal
flow.
exit
of
area,
in
terms
Exit
design
location,
and
geometry.
INTERNAL FLOW STUDIES Internal
Flow Instrumentation
The objectives investigation pressure
distributions
the engine utilizing
itudinal
stations
locations,
Kiel in
duct
in
the high
in
shown in Figure temperature and radiation
I0.
probes shield.
the
which
for pressurek long-
These rear
cylinders,
Total
The inlet
of
and above
pressure
Kiel Kiel
tubes tubes
mounted tubes
shown in Figure
are
I0 are the plenum
of a thermocouple
The temperature
the
at several
8, were at the leading
consist
Total
plenum.
9, and the cylinder Also
through
drop.
ducts.
and
techniques
pressure
line.
exit
temperature
were taken
of the
center
instrumentation
losses
pressure
in Figure
on its
located
flow
different
tubes,
front
are shown in Figure
14
baffle
illustrated
each cylinder were also
flow
and pressure
and to evaluate
surveys,
the inlet
internal
were to measure the
installation, measuring
of the
Investigation
probe
sensor
locations
are
given
in Figure
8.
The pressure the baffle
distribution
pressure
drop across
a number of different techniques
of both
included.
Figure
rations
the baffle
ternally
chamfered
increased
30 degrees.
cylinder
line
located
local
exposed
fins. to the
losses. sist
The "baffle
of a brass
intercylinder drilled
engine
for,
to pressure
fin
button"
probes
roJndheadmachine
baffle and fitted measuring
tubes
I0,
with the
without (I),
(5) fin
exhaust (5) were top of
tubes
were
passage
Figure
a 1.6 mm tube
and
(4) were
the
ll(c),
screw inserted
instrumentation.
(2)
are
on the
head tubes
at the base of barrels. with,
This
probes
on the
flush
pressure
was in-
head tubes
height
to Figure
ll(a),
to approximate-
vertically
Cylinder
cylinders,
face
configu-
angle.
barrel
Cylinder
Referring
opening
of these
(3) were located
adjacent
probe
insensitivity
local
were
1.6 mm diameter
included
The cylinder
of the cylinder.
between
various
positions
of the cylinder.
side
were measured by
manufacturers
are
and
Representative
The tube
angularity
face
shown in Figure
(I),
tubes.
9.5 mm below the
located the
probe
ll(b).
head tubes
center
probes
The vertical
shown in Figure
the
to a 60 degree
the probe
engine
and methods. and engine
All
pressure
upper
the engine
illustrates
button
total
stack
airframe II
open-end
]y
probes
and locations.
except
on the
con-
through
the
The screw is for
connection
The head of the
screw
15
is
filled
and smoothed.
were mounted
in
the upper
integrated
or averaged
The upper
plenum static
element
pressure
were attached spacing
to the
between
probe
use total
pressure
figurations
probes
der head upper
together
to give
a single
aforedescribed
16
Hole
they
probes
probe
here
is
to
are shielded probe
con-
and ll(f).
A
in
the lower
plenum
(I)
in Figure
ll(a).
to each of the (5).
All
lower
cylinplenum
were manifolded
measurement
for
configuration
that
used was the
piccolo.
in the
The pressure tape
the belts
ll(c)
adjacent
averaged
temperature
plenum are shown in Figure
analog
ll(e),
same configuration,
The fourth
The thermocouple
positioned
by multi-
The total
positions
plenum pressure
configuration.
measured
of the cowl.
so that
was located
were located
of the
pressure.
Common practice
button
probes,
an
plenum was measured by four
shown in Figures
pressure
to provide
was 5 cm.
velocities.
probes
8 and ll(d),
static
surface
located
flow
at each of the baffle
of the
upper
lower
tubes
set of baffle-shield
plenums
was also
configurations.
high
Figures
As shown in Figure
the
used are
Fin-shield
pressure
elements
in
different
any local
measurement
inside
belt
tubes,
and lower
belts.
The pressure
from
Piccolo
lower
8.
probe
Two additional
plenum,
in the
probes
upper
were
one at each exit.
and temperature
recorder
locations
usin F a serial
data were recorded multiplexing
on an
format.
A
total
of
144 channels
temperature
of pressure
data
data were available.
system was also
used for
and 48 channels
An 80-tube
additional
of
photomanometer
pressure
data when
required. The purpose probes
and techniques
plenum and lower involved
in
is
is
large,
then
the
of the
craft,
engine
relatively
for
is
is
pressure.
cowled
there
pressure.
upper
problem
of whether
If
the plenum air
same.
one
flow
If,
rates
however,
one
and the plenum volume
will
The results
face,
of cooling
are the
be a distinct For
plenum was large
small.
engine
the question
the range
tightly
and total
lower
the
different
A fundamental
or a total
small,
static the
measuring
two pressures
correspondingly
between
for
was to evaluate
plenum pressures.
a static
encountered, side
study
these measurements
is measuring volume
of this
the PA-41P test
and the upper
should
difference air-
plenum was
be interpreted
accord-
of the different
lower
ingly. Figure
12 presents
plenum pressure airspeeds,
in Figure
measurements.
altitudes,
are referenced
and cowl
to free-stream
12, all
The fin-shield
methods
probes
the baffle-shield-down baffle-shield-up
a comparison
probes
give
The data represent flap total
give
probes
settings.
All
pressure.
As indicated
essentially
a reading give
read the
different
the
3% below a reading
pressures
same measure. the piccolo
2% above.
same as the piccolo.
and The
The
17
differences i.e.,
are felt
effects
to be due to position
associated
with
the open end of the probe. the piccolo the
lower
tube appears
However,
cooling
flow of
ence
between
static
data
from
in
Figure
pressure
13
for
referenced
to
the
of
the
rear
left
is
13.
position velocities
18
be given
the to
pressure
error
effects.
of
irregular
is
the
velocities
and
in the
inlets.
to right
to
down can
than
of
be
made
regarding
of
methods
This
to
in
presented
are
having
the
given
in
static coordinate intersection
cylinders,
measurement
Pressure
progressing
bank
due
are
ordinate is
differ-
region.
free-stream
right
the
the a
data
left
directions
in
belts
the
is
with
longitudinal
The
bank
the
The
the
the
combined
correspondingly
number.
engine;
down
this
referenced
cylinder
the
and
represents
smaller
area of the upper
flow
probes
form
observations All
be shown, if
measurably
pressure
different
progressing
Several Figure
the
total
face
two
in
There
and
abscissa
front
towards the
m/sec.
coefficient The
is
15
results
engine
pressure.
of simplicity,
should
plenum is
650 cm2 ,
rate
neighborhood
the
of
method of measuring
as will
The cross-sectional
approximately
air
and orientation
consideration
of the upper
plenum.
plenum is
effects,
probes.
The volume lower
type
standpoint
to be the best
plenum volume were small,
the
location
From the
plenum pressure.
the baffle-shield-up
the
error
and
to
cylinders. the were
data
subject
finite
plenum.
The
flow
in to
increased
scatter
due to the flow.
The scatter
Since
is
tested.
of the flow
the pressure
variations,
it
is
also
plenum dynamic from
front
engine
varies
Again,
this
belt
that
pressure. is
lower
the
The increase
plenum.
recovery
in
with
static
flow
The left
side
of the plenum
from
the right
1-3-5).
This
asymmetrical
behavior
measurements
and is
is
visible
The baffle measure
of the
indicates
low.
may be raised so that flow
it
button engine It
is
probes face
button
(I)
extend
provide
that the
through
side
believed
to be
governor.
The 9.
the most reliable
The piccolo
tube
the piccolo
reading
tube by cutting
it
to the front
the
showed itself
shown in Figure
pressure.
biasing
cylinders
short
where
the
highest.
In summary, results baffle
the inlet
is possible
through
does not
velocity
by the propeller
inside
effect
of the
(cylinders
governor
the
passage
differently
blockage
of
pressure
the diffusion
behaves
flow
pressure
at the engine
2-4-6)
due to inlet
a change in
static
(cylinders
in a number of different
to the
the plenum.
the pressure
consistent
to be
slipstream
indicates
is no effective
the progressive
to the
is believed
according
measures
There
to rear
accompanying
also
through
evident
static.
condition
change in the propeller
pattern
inlets
character
face
the climb
corresponding
different the
for
probe
gives
of this
investigation
the most reliable
show that measure
the
of engine
19
face
pressure.
cylinder
baffles,
equally
well.
accurately small,
If
the engine then
then
not
a shortened
The pressure
by the
is
equipped
piccolo
in the lower
piccolo.
However,
the baffle-static-up
with
tube will
work
plenum is
if
this
probe
inter-
measured
volume
is
configuration
should
be considered.
Engine Background. cooling the
air
similar
to
Engine
mass
pressure
Orifice
flow
drop
across
orifice
flow,
general
power
In
A
tation
(I),
study law
the
the
engine's
engine.
which
is
The
fin
relate
this
w
= a
to
relationship
described
the
passages
is
by
= a p_-_p.
of
of
of
"b"
has
(I)
relationship
is
given
engine
(pap)b
the
air-cooled
relationship
proportionality
2O
characteristics
by
the
law
equation
power
form
orifice
through
w
A more
Characteristics
of
orifice "a"
(2)
value
engine
of
(2)
characteristics. as
=
0.5.
development
equation
functions
b
an
is
shows
a valid The
equivalent
that
the
represen-
constant
of
orifice
area
in
that
it
and total fin
tends
to vary
passage
area.
spacing
reported
in reference
current
ranging
with
to b = 0.50
for
range,
is
depending
spacing
of tests
from b = 0.78
aircraft
to b = 0.58.
spacing
a function
Fin
ranging
fin
5 mm spacing.
horizontally-opposed
a much wider
"b"
baffling.
13 show "b"
from b = 0.52
cooling
The exponent
and intercylinder
0.5 mm spacing for
directly
Available
engines
Values
on geometric
for data
has "b"
of "a"
vary
over
engine
size
and
number of cylinders. The values
of "a"
are determined that
by ground
shown in Figure
sponding
baffle
In
applying
linear
in
through
the
because
aircraft
the
drops
engine
direct
corresponding
theory, in
cooling engine
and "b"
in
the range
are then
deter-
form,
(a)
coefficients
(3)
measure This
can the
mass
baffle
be
do flow
pressure
used
cooling
approach
installations air
to
and corre-
over
logarithmic
model
techniques.
flight.
engine
mass flows
(_Ap) + In
to
engine
a system similar
are measured "a"
regression the
a particular
air
(2)
(w) = b " In
(2),
to
Cooling
equation
determined,
equation
selves
7.
for using
The coefficients
mined by rewriting
Once
test,
pressure
of interest.
and
and "b"
not
air
flow
is necessary lend
measurements, drop
with
is
themwhereas readily
21
measureable. ments
For
then,
the
validity
of
altitude
and
mined
engine
drop
with
in
a
Altitude
research
ground al
maries are
of
given
and
engine
term
in
An
density into
22
ratio
"a".
orifice
The
14
work
test
flow
the
deterthe for
ground
will
be
will
and
the
test
discussed be
dealt
and
was
which absorb
orifice
the
is
extrapolate to
operationSum-
which
the
is
level now
has
cooling
exit
to
flow
density
stagnation
made
equation
to
settings.
method
This
sea
correlation
investigations
(3) with
and
charac-
engine
data
entering
cylinders.
modification
engine
cooling
mixture
The
the
pressure
and
and
orifice
air-cooled
effort
15.
static
"_"
the
characteristics
(2) the
engine
requirements
replaces
leaving
of
this
power
equations
flow
additional
The
between
second
radial
cooling
references
density
the
for
on
part
of
different
this
on
and
influences
objective
determined
from
based
heating
The
air
ground
cooling
the
with
heated
the
while
associated
in
of
measure-
considerations:
correlation
concerns
effort
the
the
flow
meter.
two
these
as
emerged
the
of
investigated
altitudes
on
and
were
test
orifice on
configuration first
cooling
section. and
development.
an
effects
paragraphs,
later
teristics
as
measurements
The
followin_
flight
depend
heating
installation
the
used
characteristics;
configuration. in
is
of
measurements
pressure
aircraft
purpose
engine
such
orifice
baffle
the
density
of is
temperature. utilize
density becomes
the
constant
w = a(Oexa p )b.
Equation 12,000
(4) has been shown to work meters
in
altitude.
results
of reference
polation
capabilities
density
ratio
of the
is
increasing
with
higher
Use of the ambient
power law,
on the
unaffected
by altitude,
Therefore,
is
a well
proven
characteristics,
for
relationship,
flight
Flight
test
test
solely
should
would
density
a valid
ratio
extrapolation,
mass flows.
engine
of heating orifice
relationship
derived
be-
and altitudes.
influence
the necessary
This
which
The exit
determined
for
the power law
parameter
to the
test
to provide
use of this data
in regard
a
and exit
and altitude.
velocities
up to higher
extra-
works
effects
provides
the
versus
parameter
deviations. hand,
ratio
However,
ratio
from
the ambient
ratio
flow
other
on ground there
orifice
of
mass flows
density
in even greater
available
density
up to
of the
density
caused by compressibility
come dominant
istics,
exit
and low altitudes.
down with
breakdown
altitude
shows a comparison
The average
low mass flows
result
14,
effectively 14, taken
based on the average
densities.
breaks
Figure
(4)
character-
[equation
extrapolation. from ground be considered
and
(4)]
Altitude
test
data
and
as valid
use.
measurements.
The
flight
cooling
air
mass
23
_
flow
measurement
system
consists
of an array
mounted
in each inlet
azimuthally designs
in the
internal suited
across these
of flow
fitted
involving nomial y
= 0
static duct,
also
24
to
the
products because
and the
and
+
probe
data
static
effects
observed
of the
form
(5)
fitted
data.
usable
involved
in
polynomial across
mass
Terms the
either
flow
the rate
on
of
the
x for
Installation
sides
data.
+ a 5 x 3 + a6y3
not
air
in the
integration. both
re-
pressure
surfaces
distributions cooling
the
No indica-
pressure
y were
the
of the
that
with
was observed
location
con-
studies
wake influences.
and
pressure
the
showed some variation;
a 3 x 2 + a 4 y2
Using
numerical
pressure
are axisymmetric
and total
technique,
x and
corresponding
recorded.
duct
or stall
total of
all
total
by
inlet
distributed
indicated
were consistent
a2v_
coordinates.
determined ature
the
least-squares
+
ports
to form
The static
seDaration
P = a1z
probes
Previous
inlets
and propeller
the
pressure
behaved and no adverse
variations
of attack
Using
were
configuration. of these
The system
The inlets
the nose cowl
from the propeller.
however,
tion
into
15.
pressure
ducts.
was well
distributions
angle
inlet
behavior
flow
and static
and static
"bug eye"
aerodynamic
shown in Figure
of total
incorporated
ventional
is
engine
poly= 0 the
inlet was temperwere
or
Data 1,800
were
m
used
to
to
three
7,200
vary
is
total
and
this
the
effective
Equation
(6)
logrithmic the
with
onto
ordinate
altitude.
curve
Figure
a
used
a
ratio
(Ap)
+
In
a family
of The
which,
Using
corrected
the
'', the level
separated curve
as
as
were
Figure
16.
The
and
subsequent
in
upper The
by it
"a",
one
(a
Oex
plenum
reason
The
placing
coefficient
from
from
predicted and
m
data
section.
behavior
In
the
pressures. later
900
settings
between
intercept
sea
in
of
flap of
here,
coordinates.
"_exAP
single
the
represents
scaled
parameters,
in
= b
sample
given
static
area
cowl
difference
density
orifice
(w)
is
the
with
A
value
discussed
exit
In
is
as
increments
and
flow.
plenum
consistent
Taking
mass
taken
be
Airspeed
drop
lower
will
altitude
altitudes
pressure
graphs,
in
m.
the
different
baffle
are
taken
results
equation with
b)
the
(6)
.
lines
last
term
shown
in
on
in the
Figure
baffle
pressure
altitude
curves
in
(4).
has
straight
shown
for
Figure
right 16,
varies
drop collapse 17.
The
follows
18
In
(w)
presents
a
= b
" In
comparison
(OexA p)
between
+
In
the
(a)
result
of
Figure
25
17 and manufacturer's test
aircraft.
with
the respective
The "a"
data
The difference
of the
values
installed
in the
aircraft
The implication
the engine
baffle
system.
below
that
the
significant
reduction,
and unheated
air
entering
lending
was by-
the external
measurements
the cooling
air,
is test
the intakes
through
flow
and suggested
cooling
areas.
drawn was that approximately
and leaking
and at
test
lies
of the manufacturer's
Flow temperature
engine
of the
two curves
engine
than
of the cooling
the
orifice
cell
passing
between
the effective
larger
engine.
to the engines
of
significantly
55 percent
applicable
exit
made directly showed a
the mixing
support
engine
of heated
to the
leakage
theory. The external Figure tape
19, is
pressure high
is
of the occur
typical
sides
of side
as it
To test placed
26
is
is
forced
however,
the leakage
by a cover
of the
from
against
rubber
and low
a view
looking
seen laying
front
the
inside
to do.
Sealing
is
tighter
against
the cowl by the
flow.
A simple
apparently theory,
on top of
the high
20 is
engine
shown in
A neoprene
between
Figure
is
intended
of the cooling
method in theory,
seal
tape
the engine,
practice.
of the cowled
when the tape
ram pressure
the
the engine.
The neoprene cowl
system about
of current
used to provide
pressure
to rear.
baffle
not
and effective so in practice.
the neoprene
the engine
assumed to
tape was re-
as shown in Figure
21.
This in
"do_ house"
this
passive
modification
region,
and removed
systems
such as the rubber
were reoeated
and the
"dog house"
reduced
the leakage
about
engine
baffle
cooling
flow
discrepancy testing.
improved
at this
To attack
in reference
Ground
test can
as
ground
flight,
be
particularly
test
the
flow
internal
air
well
by
a blower
on
If
case
when
Flight
the
an
same less
sufficient additional
of the validity flight
installation
configuration
question,
an
should
toward
ting
the
38 percent
to warrant
the
cell
testing
are
however,
decreased
indicated
curve
performed
aerodynamics
points,
still
validly
system
two
The
a test
came to
technique
4 was employed.
aerodynamics.
external
i.e.,
with
between
this
22.
"dog house",
the question
measurements.
investigation
internal
point,
tests
inlets.
curve
and the ground
the forefront.
is
the
the measurements
configuration
This
through
from the manufacturer's
of comparing
utilized
the
drop is obtained
enterin_
Also,
The flight
area,
seal
accompanies
in Figure
the system orifice With
lock
that
tape.
are given
the engine.
this
a positive
the uncertainty
results
pressure
However,
in
provided
at
flight flow, the
the
ground.
related
the
ground
strongly
the
investigation
the
inlet only could Ground
as
concern
and
involves
be
of is
exit.
as
cooling
directed
Between
a means accomplished
test
well
considered.
investigations
if
serves which
on be
test
required
aircraft
of
these
generajust
investigations
as of
27
the
internal
advantages flow
aerodynamics of unconstrained
measurements,
vantage
finger
offer
variations
most importantly,
observation
A moistened
installations
configurational
and perhaps
of personal
system.
of cooling
of the
or
the and
the ad-
functioning
cheek are very
of the
effective
leak
detectors. The ground
test
system consists flow
system is
of a variable
metering
section,
shown in Figure speed axial
diffuser,
flow
23.
The
fan,
mass
and connecting
ductwork
as required. Tbe first
configuration
shown in Figure the
f!i_ht
parisons section
mass flow between
During valve,
this lifter
and engine
Ing.
leakage With
primarily
system
in the
inlet
and the
rakes
through
and the
front
the prop
external
by this
engine
with
about
and out
and nose cowl.
the
baffle duct
tape
A reduction
additional
was still
baffle
The
correction.
was observed
leakage
spinner
this
was repeated.
was obtained
Com-
correction
results.
were sealed
test
configuration, the
test
the metal
These regions
8 percent
this
leakage
inlets.
This
18, and 22 include
and between
rubber, of
to the flight
was to validate
duct metering
of 17 percent.
additional
proper.
the gap between
28
measurement
16, 17,
covers
and silicone in
test
the
test,
"dog house"
of this
retroactively
in Figures
was the
The purpose
showed a discrepancy
was applied data
21.
tested
seal-
detectable, through The presence
of the nose cowl made it accordingly, installed This
it
arrangement,
the high
no leakage,
flight
was important
configuration
"ideal
cell
engine
case"
flight
baffle
ductwork
24, represents
it
served
comparison
in Figure
configuration
the maxThis
as the flight
with
the
ground
configurations
25.
Testing
showed that for
was engine.
case.
These two test
was responsible
region;
plenum of the
that
for
schematically
maximum seal
this
configuration in
configuration.
are presented
pressure
shown in Figure
configuration
test
to seal
was removed and additional
to reach
imum seal,
impossible
the
an additional
of the front
9 percent
leakage. Referring figuration of flow
to Figure
represents behavior.
of the engine as it
will
pressure baffle
ever, air
pressure
enters
oil
about coolers,
large
case"
enters
test from
the high
manner and in the
the cooling
fin
cell the
constandpoint
pressure
passages.
and the velocity
is
plenum
case of the
static
and ambient
flight
configuration,
at much higher to the obstacles
fin
velocities passages.
directly
prop governors,
in its intake
side
same direction
low.
drop is measured as the difference
pressure
in the
"ideal
in a uniform
move through
perpendicular flow
the The flow
plenum is
the high
25, the ground
The high The between
static. the
Howcooling
and in a direction In many cases, path
it
must
such as alternators,
manifolds,
etc.
A
29
significant
wake in the
due to
the "bug eye"
static
pressure
intake
Total
particularly
with
concern
is whether
determined to those vestigate
this
was assembled
the
engine.
A total
a uniform in the
results test. both
drop for
the
plenum total, Use of the cent. rear 30
orifice
flight rather
The location engine
baffle.
uniform
pressure cell
26.
"ideal"
flight
total
if
gives pressure
Data presented
confirmed
The rewith
the
configuration as measured by the baffle
pressure
pressure
at the
drop was measured
is based on high
than on high
of the
engine
27 alon_
characteristics
configuration
flow
practice.
in Figure
are identical
plenum static
To in-
shown in Figure
test
are given
configur,qtions
comparable
configuration
made at the
baffle
from the maximum seal
test
cell
an "ideal"
survey
The engine
The engine
are
in the bend and honeycomb in
to achieve
test
of
performed.
is
same manner as standard from this
test
cir-
characteristics
measurements.
measurements
configuration
these
The Doint
measurements test
flow
essy to acquire,
orifice
a ground
pressure
flow.
engine
by flight
cell
duct
are relativel.y
cell
vanes were utilized
vertical
sults
test
and flow
The test Turning
the
question,
immediately,
A true
use of Kie_ probes.
by ground determined
develops
to measure under
pressures the
flow
configuration.
is difficult
cumstances.
here
internal
pressure pressure
plenum static.
an error probes
in a later
of 8 perwas at the section
show
this
region
to be the
].east
affected
and blockage.
The key finding
is
test
that
istics
ground
cell
are perfectly
the flight
rather
orifice
determined
valid
measurements
pressure
than
pressure
for
this
particular
engine
flight
configuration
static.
test
and lower
plenum
drop parameter.
This
all
engine utilize
to form
includes
if
plenum total
herein
static
character-
utilization
Accordingly,
data presented
test
orifice
are based on engine
characteristic
plenum total
from
by inlet
upper
the baffle
Figures
16,
17,
between
the
18, and 22. In order flight
to achieve
and ground
introduced data.
flight
test
istics
of both
settings.
to ambient
exit
density
However, curve
"hot"
if
the
engine
temperature,
Therefore, a correlation
orifice
ratio
is
the orifice
shut
the
appears
that
factor,
works well
the
ratio
are
data power to
data. the
same
altitude
correction.
is
the
used,
"hot"
as shown in Figure
use of the for
cruise
"cold"
characteristics
density
character-
down and allowed
to obtain
used for
a
The "hot"
at normal
downward by I0 percent it
correction,
engine.
operating
was then
the ambient
shifts
and "cold"
must be
data to "hot"
this
to measure
the
and "cold"
engine
to determine
The engine
The "hot" if
In order
with
a correction
the "cold"
was performed
was taken
comparison
results,
to convert
engine
cool
test
a valid
exit both
density altitude
ratio
28. as
correction, 31
and "hot"
and "cold"
The results rized
engine
from the
in Figure
29.
It
flight
is
significant
leakage
problem
rubber
external
baffle
tape
that
the
test
that
the
external
aircraft
can visualize
two adverse cooling since
stream less
Reducing through
leakage the
engine Cooling
with
Leakage produces
proDortional
air
through
and the
passes
total
through
produce
drag and reduced
the
reduction
the
losses air
in
the baffle
and m!enum.
mass flow, the
due
The
immediately
an increase
Correlation
The
the engine,
or around
and, correspondingly,
to
installation.
across
inlet
one the
directly
drag is
One of the many technological
32
good shade,
is more obvious
of the
will
accordingly,
may exist
in service.
between
inlets
it
which
can be generated
are functions
of whether
in relatively
from the
losses
and,
fact
of these
results
of the
low time
the
a
The first flow
drop which
Considering
that
cooling
second
operation
the use of the
system.
figure
increased
of cooling
to increased
with
aircraft
the mass flow
pressure
exists
problem
internal
effect
the
leakage
effects:
are summa-
from
are
time
tests
evident
has very
performance. the
and ground
baffles
the
number of high
correction.
downreFard-
engine.
in cooling imDroved
flow
cooling.
Investigation outgrowths
of the
radial
air-cooled
engine
Analysis
development
was the NACA Cooling
This
is well
procedure.
procedure
numerous NACA reports
(see references
pose of the method is
to take
cooling
requirements
operational velopment
of empirical
generated
by the
ground
test.
which
investigations
was always
test
This
however, the
during
World
War II.
much reduced
liability and engine
ground
test
A flight whether could
cell
data,
test
an installed be developed
problems.
de-
the
the heat
in the
aviation
cooperation
to assist
in
the
is
with
today
product airframe
of installation
cool-
in many instances, armed only
by Figure
with
5.
to investigate
correlation solution
was a cooling
manufacturers
between
problems
cooling
between
industry.
in conjunction
program was performed engine
significance,
manufacturer,
cooling
flight
cooperation
solution
by
of ground
and engine
as represented
The objective
required
cooperation
general
work against
to solve
to
availability
technological
situation
The airframe
must work alone
requirements
be determined
to be of little
to the
manufacturers
problems.
to the
technological
This
market
concerns
only
manufacturers
in regard
The competitive
in_
tied
proved
airframe
of engine
concerning
could
in
The pur-
use of the method for
due to the close
government,
these
relationships
Consequently,
results.
ground based data
The NACA procedure
engine,
documented
16 and 17).
and extrapolate
altitudes.
Correlation
procedure of cooling
correlation
relation 33
which in
involved
flight.
which
could
be easily
measured
The NACA method was used as a basis
reference is
parameters
16).
transferred
The heat
generated
to the cylinders
H = ci
(see
by combustion
is
given
which
by
WI "c (Tg - Th)
(8)
T!
The
air
charge
indicated
flow
engine
"W c
power
is "I",
directly and
I TM
H
The
heat
given
up
=
c2
by
(8)
can
be
to
the
rewritten
as
(9)
_
(Tg
the
relatable
T h) .
cylinders
to
the
cooling
air
flow
is
k H
For
flight
tical are is
to used
= c3
testing,
the
measure. here
The
and
the
(T h
- Ta).
cooling
air
orifice
(I0)
mass
flow
characteristics
corrected
baffle
pressure
"w" of
is the
drop
For (II)
a constant must
be
c4
cylinder equal,
(OexAP)
head
giving
n
(T h
engine parameter
(ii)
- T a)
temperature,
equations
_
imprac-
substituted.
H =
34
w
(9)
and
_
(Th
Equation
(12)
relation.
determined
had
to
be
equation
are
For
the
PA-41P
test
let
temperature
was
term
test
m.
At
each
altitude,
three
mixture
settings
ranged
supercharging,
the
test
cooling
Altitude
from the
a
same
altitudes. air
flow
test
results
are
four
each
of
to
was peak
power cowl
run. EGT.
temperature.
were
in-
becomes
(13)
n
orifice
1,800 four The Due
settings flaps
test
of
to
turbine
the
from
in
decided
gas
now
matrix
settings
The at
part
ranged
full-rich
was
= cIm/(cexAP)
as
"
generating
quantities
it
(12)
g
back-pressure,
supercharger
= T*
flown
heat
exhaust
Equation
was
other
flight,
"T
relationships
physical
the
- T h)
temperature
exhaust
measured
used.
T
and
other in
the
study.
and
all
all
Correlation
Empirical
ratio,
aircraft,
program
settings
at
with
- Ta)/(TEG
characteristics 7,200
As
gas
it
measurable
this
The
between fuel/air
replace
(T h
Cooling
NACA
testing.
timing.
(12)
the
by
as
(12)
p
combustion
ground
)n
clm/(Oex&
effective
such
ignition
Th)
essentially
established
parameters and
is
The
was
_
T a) / (Tg
m
to
power mixture to
turbo-
were
obtainable
used
to
vary
point.
Q
The Figure
30,
the
presented
relationships
in
Figures
between
engine
30
and baffle
31.
In
pressure
35
drop and the of
temperature
the cylinder
stant
indicated
altitudes
are plotted
power.
represents
but
equation
with
(OexAP) as the
(13).
in(T*)
The resulting
cooling
31.
hottest
running
of the
to those
averaged
of Figures
relations
are given
in
four
of
curves
intercepts,
in agreement
logrithmic
indicated
form
(14)
engine
relation
correlation
cylinder
a range
+ In(cl m)
the
cylinder
is
a con-
variable,
correlation
A separate
for
ordinated
(13)
= -n'In(oexAP)
with
All
This
independent
varies
place
different
Rewriting
The intercept
the
mass flows.
on the power setting.
with
Figure
Each plot air
same slope,
dependent
T_, based on the average
head temperatures
and cooling
have the
ratio,
is
power I. given
was performed
temperature, temperature.
30 and 31 were obtained.
in
using
(cylinder
number 6),
Results
similar
Both cor-
below.
TA* = (ThA- Ta)/(TEG T _ ThA ) = 0.1710"52/(OexAP)
0 29 •
T6* = (Th6 - Ta)/(TEG T - Th6 ) = 0.1810.50/(aexAp)0.28
Both
show excellent
by equation 36
(13).
in
agreement
with
the behavior
predicted
(15)
(16)
The results correlation
for
developed able
of this a particular
in terms
in flight.
provides
This
and for
relating
mance.
The airframe
the
cooling
on _round
stallation
flight
test
Internal
tions
exit.
air
for
cooling, leads
etc. to the
installation
using
Appropriate the heat Figures
for of
the
locations
which
the basic
32 and 33.
cooling
in-
Rise
correla-
rise
is
important
in that
of the airflow
is is
of the
plenum and at the
such as oil
analyses.
temperature
from
analyses,
the upper
cooling,
density
transfer
for
perfor-
freed
temperature
The second location exit
thus
data,
use a portion
design
developed
is
correlation
of these
auxiliary
problems,
to aircraft
Flow Temperature
at the rear
The first
measure-
investigations.
developed
many installations point
cell
are easily
of cooling
requirements
of the coolin_
were also
cooling
solution
test
can be
once established,
manufacturer
dependency
a cooling
installation
which
correlation,
for
show that
aircraft
of quanitities
the basis
As part
investigation
at this
cooling,
inter-
of interest
since
important
for
The correlations
formulation
it
future were
of equation
(13).
terms were changed to represent
process. For the
The results range
are given
of baffle
in
pressure
drop 37
(cooling craft,
air
mass flow)
the cooling
flow
developed
by the PA-41P test
temperature
rise
at the
air-
rear
baffle
is 20°C < (Tup - Ta)
and the
temperature
rise
across
< 30 ° ,
(17)
the engine
is
(18)
70°C < (Tex - Ta) < 100°C.
The exit
temperature
numbers reported engines. should here
The upper
on the flow and these,
in
are
values,
mechanism
cooling.
to the cooling turbulence
dependent
however,
transfer
to as velocity
and flow
with
horizontally-opposed
The heat
transferred
turn,
for
plenum temperature
carefully.
velocity
are consistent
sources
commonly referred
amount of heat
engine
values
by other
be treated is
rise
air
The
is
dependent
in the plenum,
on installation
and
configurations.
INLET INVESTIGATION Background The function dynamic pressure pressure 38
of the
inlet
and deliver
plenum in a uniform
is
to recover
the cooling manner.
flow
Ideally,
the available to the high this
should
be accomplished tion.
with
Inlets
as in the
are classified
case of wing
dimensional, inlets
no internal
as in
in comparison
turbojet
powered aircraft.
configuration allows
for
inlet
with
air
is
aerodynamic
analysis
are the absence
is
of practical
As part
this
effects
testing
The design
on the
and the
used.
of
through
the a
little
The reasons methods
of different
for and
candidate
a systematic cooling
one
Very
analytical
program,
area
through
and intuition.
and design
shapes.
and their
frontal
which
to use two inlets,
spinner.
dictates
of aerodynamic of
is
and to the
engine
to pass vertically
air-
complex
turboprop
has been accomplished
the cost
The
engine
due primarily
minimal
practice
shape heretofore of styling
inlets
used for
This
of the propeller
combination
this
reciprocating
to those
Conventional
on eacb side
or three-
and are of relatively
an installation cooling
separa-
intakes.
of the horizontally-opposed
requires
engine.
edge intakes,
aviation
are three-dimensional,
flow
two-dimensional,
the case of axisymmetric
geometry
yet
as either
leading
used on general
craft
or external
study
installation
of was
performed. Inlet The theoretical be discussed
Aerodynamic aerodynamic
in relation
Theory behavior
to three-dimensional
of inlets
will
axisymmetric
39
configurations.
Figure
pertinent
parameters
design
of the inlet
is
and recover in static let i.e.,
is
velocity. small
A large
through If
internal
and vice
internal
diffuser
length
is
also
is
free
corresponds
stream to a
The diffusion, takes
place
available,
possible
in-
(Vi/Vo),
to the
versa.
occurs,
of the
ratio
recovery
the recovery
sufficient
recovery
velocity
pressure
ratio
which
inlet
amount of flow
head as an increase
to the velocity
of the
velocity
dynamic
the
The purpose
required
The desired
related
the ratio
the
of the
pressure.
with
identified.
to capture
a part
itself
34 shows an inlet
externally.
then an
to increase
the
re-
covery. The ability
of an inlet
pressure
recovery
and,
properly
over a range
in fact,
to its
cross-sectional
Figure
34.
The axisymmetric
are from
reference
18.
segments
which
The el]i_sesare designation.
the desired
ability
inlets
used in
at the nose of the
proportions.
in Figure
35.
inlet
re-
according
families
part
the inves-
lip
in
elliptical contour.
to the
"A"
are available
Three A-series
The numerical
in
detailed
of two distinct
in proportion Other
is
shape as illustrated
They consist join
to function
conditions,
the KuchemannA-series,
different
4O
its
of operating
lated
tigation
to deliver
inlets
for are
of the designation
shown
refers
to the percentage
maximum external produce large
thick
In Figure
distributions different
velocity
19.
adverse
of the
gradient
of the lip curvature stagnation
employed.
peak is
is
small
velocity
large
layer for
and external
flow
there
pressure is
is
of the
inside,
the radii
thickness
on the
radii
ratios,
of the
lip
nose,
of curvature.
suction
peak correspondingly therefore,
inlets
and small
external
point
of
the
due to the
stagnation
are
and stall.
on the relative
outside
a
included,
peak and severity
velocity
to
gradient,
separation
the
ratios
pres-
similar
response
on the
with
The
to the method
correspondingly,
to the
for
of attack.
ratios,
In _eneral, velocity
internal
For larger
formed
in combination
and the
are given
according
the
pressure
characteristics
suction
or,
angle
outside.
contour
are dependent
contour
point
flow
by an adverse
conditions
relatively
numbers produce
For many conditions,
when the boundary
The strength
with
and angles
exhibit
peak followed
the necessary
inside
lip
Both the
of airfoils.
which,
suction
the
area to
Low numbers
contours
and high
ratios
sure distributions
suction
area.
were calculated
of reference
inlet
36, the potential
about
distributions
those
lip
of curvature
reverse.
of the
cross-sectional
relatively radii
ratio
lies forms
designed pressure
the
turning For to the to the for recovery
41
are
subject
tion
to the possibility
and stall.
ratios
Inlets
and large
wise
subject
of internal
designed
external
for
pressure
to the possibility
flow
small
separa-
velocity
recovery
are like-
of external
flow
sep-
aration. The effect on inlet
pressure
design tion the
tenet
flow
inlet
inlet
lip
large
of curvature
peaks and following
contours
Five
different
in this
program.
consist
of the
locations
inlet
should
of the Test
PA-41P inlet,
current
aviation
of reducing The propeller
42
the
as STD, is propeller
practice.
hub incorporates
with
suction
gradients. Inlets were investigated 37-41.
three design
They
axisymmetric representative
The PA-41P inlet,
a swept configuration, blockage
stagnation
ill-defined.
They are shown in Figures original
shape is not
be well-rounded
configurations
and a conventional
designated
of
technology,
inadvertent
pressure
configurations, general
the reciprocacurrent
are also
to minimize
adverse Design
for
nose cowl
contour
Accordingly, radii
the
point
one fundamental
With
an arbitrary
lip
stagnation
suggests
be followed
Therefore,
on the
of the
installations.
about
defined.
points
should
cooling
field
location
distributions
which
engine
well
of the
by relieving a 20 cm shaft
with the
of here
the purpose frontal
extension
area. to
provide
the necessary
effects
of sweep for
The three the
true
according (0.3
for
to their
incorporating
is
for
this
The STD, 0.3F, one program
effects
of the pressure
and the baffle The baffle air
of
general
0.3F inlet is
these
ratio or aft).
aviation
design
shape as possible.
GAC. Effects
phase,
baffling,
were investigated with
the
some modifica-
0.3F and GAC
were tested. The internal
terms
to
a point
(forward
and 0.3A inlets
engine
of
are designated
velocity
location
In a later
to the external
inlets
design
Internal
phase.
provides
The inlets
inlet
0.6F,
applies
of incorporating
a conventional
Inlet
tions
it
the results
as much of the
The designation
in
information
respective
because
data base and analytical
this
and longitudinal
inlet
are unknown.
were investigated
configurations,
comparing
The aerodynamic
of inlets
a cowl nose piece.
or 0.6)
The fifth
this.
types
inlets
While
axisymmetric
shapes into
for
of an experimental
procedures.
reference
these
axisymmetric
existence
design
length
mass flow,
indication
recovery
pressure
pressure
of the
as seen in Figure
are measurable
obtained
drop developed
drop varies
of the volume
inlets
in
the upper
plenum
across
the
engine.
with
the
cooling
directly
5, and therefore
of cooling
in
flow
obtained.
is
an The
43 \
pressure
recovery
represents the
one
other
or
available
for
drag
for
ler
the
same
in
disk
as
coefficient
in
are
given
a planform
the
view
is
seen
available
spinner
cruise,
and
Dressure
to
near
the
with
condition,
i.e.,
operating
occurs
for Figures
pressure
44
inlet
43-46
surveys
the
the
results
inlets,
inlet was
to
free
for
the
for
shank
four
cooling
the
nature
of
of
42
in
This
The
are
attributed shanks of
On
the
in
cruise.
the
varies
for test
the
this pressure
the
of
total 0.85q_
outer
edge
in
for
total
losses
operating
average,
from
of
reductions
attack
propel-
from
the
to
A
the
the
effect
at
flow
Superimposed
distribution
climb.
results
of
static.
l.lq=
angle
to
terms
the
defined.
results
The
the
be
close
stream
inlet.
of
to
The
nacelle.
terms
give
engine.
higher
needs
37(a). Figure
flow
head
in
mounted
i_
condition. the
through
propeller
the
pressure
flow
the
spinner
the
system;
cooling
approximately
1.2q=
Poor
it
flow.
the
at
to
associated
the
in
drop.
the
pressure
Figure
of
in
less
rake
shown
losses
First,
means
evaluate of
reasons.
losses,
recovery
referenced
propeller
at
the
survey
shown
pressure
inlet
front
two
pressure
cooling
properly
for
internal
baffle
moving
pressure
survey
is
the
two
pressure
immediately total
important
the
large
poor
To
of
being
recovery,
Second,
is
in
is a
upper
inlets.
a
too
loss
stalled
high of
plenum The
for
0.1q=
total survey
points
were at the end of the
row of cylinders tion
(see Figures
shown is
the upper
and upper
cylinders
the graph. inlet
a top view
left
respective
in
in
lower
combination
recovery
recovery
in
its
internal
stall.
the propeller to the
left
This
inlet,
than
the right
stall
angle.
swirl
component
is
obtained
fromthe
inlets.
and consequently
The angle
of
attack
of the propeller
through internal
the
in
flow
the presence
left
a
the
ex-
in pressure of an
the blockage inlet,
angle
Also
by
or due
of attack
exceeding
asymmetry flow.
re-
area diffuser with
at a higher
similar
pressures.
associated
in the
behaved
and STD inlets
by the inlet,
indicating
pressure
distribution
43 shows a loss
located
operating
inlet
the lift
are well
may be due to either
governor inlet
0.3A,
an expanding
in Figure
left
in
plenum total
diffusion
the plenum volume
The STD inlet
The
1-6 outside
the highest
ducts
of the plenum
and the diffusion
into
in
graph.
located
pressure
of upper
by incorporating
duct,
is
44 gives
The 0.6F,
values
inlets
by the numbers
The inlet
of external
diffusion
panding
in Figure
rake.
the
number 2.
by showing a total
The pressure
inlet
of the
of cylinder
to the propeller sulted
right
corners
each
The presenta-
with
governor
the plenum.
as indicated
and across
8, 9, and I0).
are denoted
The 0.3F inlet
duct
of the engine
The propeller
in front
recovery
inlet
is
the
design's
due to the
evident
in 45
the
figure
is
the nonuniformity
plenum as indicated This
again
is
of the flow
by the variation
the result
in
entering
total
of insufficient
the
pressure.
diffusion
by the
inlet. The 0.3F inlet
in Figure
recovery
in the plenum.
dicating
that
externally
part
The flow
of the recovery
by the
in Figure
than
the
0.3F.
suit
primarily
duct
and the plenum entrance.
obstructed external inlet
from the poor
by the
front
diffusion
moves aft,
dynamics,
is
in good pressure
is more uniform,
in-
has been accomplished
inlet.
The 0.3A inlet sure recovery
44 results
45 produces This
loss
interface The inlet
cylinders.
geometry,
increasingly
is
less
believed
between duct
the is
The extent
was accomplished its
0.2q_
is
which
compromised
to reinlet
partially
to which
unknown.
controls
pres-
As the
its
aero-
by the nose cowl
geometry. The 0.6F While the
the
loss
46
the
in pressure
condition
The inlet
for
configured
conditions,
is
46 shows similar
was designed
0.6F was designed
cruise
This
in Figure
0.3F inlet
appropriately
the
inlet
in
internal internal
inlet
operating
for
at a higher
external
diffusion duct.
indicates
the region
is more severe
for
diffusion, using
For both an internal
adjacent climb
poor recovery.
to the
than for
velocity
ratio
an climb
and
stall
by
spinner.
cruise. in climb
due to the use of cowl
flaps
the
to the
system.
Referring
in Figure
36,
following
adverse
towards
this
increases
house"
modification
baffle.
Cooling
external
baffle
just
for
measurements
were operating
than anticipated.
at This
and consequently The results
is
test.
due to reducing
the
inlet
duct
interface
These losses
were reduced
as a result
ing air
by elminating
the baffle
in Figure
48 shows about
let.
The absence
inlet
duct
interpreting other
for
of the
these
increases
the results.
low performance
occurred
tests
and significant
rakes
flow
higher internal
operated
with
in Figures
improvement the flow with
over losses
the plenum.
of reducing leakage.
the coolThe GAC inlet
as the
STD in-
at the end of the the
The indication
inlets,
in-
used as a reference
the same performance survey
The four
are given
the earlier
is
had shown the
was also it
"dog
engine
aggravated
shows a slight
This
the
significantly
The 0.3F inlet
flow
after
leakage.
The 0.3F inlet
purposes.
with
peak and
the tendency
time
47 and 48.
associated
through
contours
suction
increasing
to have considerable
baffle
comparison
internal
at a later
mass flow
problems.
the modified
ratio
had been made to the external air
ratios
aerodynamic
velocity
thereby
was tested
discussed
velocity
flow
stall.
The GAC inlet
lets
inlet the
gradient,
internal
to pump additional
difficulty is
that,
little
or no external
losses
were created
of like
the
diffusion at the
47
plenum entrance. contour
Additional
pressure
information
concerning
distribution
are given
of the inlets
on cooling
the
in a later
section. The effects performance
over
are given are the
drop,
Figure
49 for
speed.
Only
for
which
the
test. this
cies
is
rate. climb
indicated
engine
baffle
directly
related
to the
cooling
at a constant inlets
altitude.
This
in
0.3F inlet, recovery There
48
i.e.,
however, which
the
is more energy
total
slipstream
result
into
in the
cooling
altitude
dependen-
maintaining
flow
this
lower
rake
with
pressure and also
flaps. same baffle mass flow.
at a higher internal
at the
the
survey
recovery
air
to a
adequate
total
in the
same cooling
air-
pressure
of the cowl
accomplished
translates
equivalent
the propeller
in plenum pressure
The 0.3F and STD inlets drop,
While
slipstream
pumping effectiveness
pressure
allow
effects.
in
subjected
The indicated
reduction
caused a reduction reduced
shown were
power in the climb, in
are presented
I00 kts
did not
condition.
a reduction
of importance
and the
are due to propeller
same engine
of aircraft
recovery
These parameters
three
range
The parameters
The 0.3A inlet
flight
operating
49-52.
plenum pressure
mass flow
climb
complete
in Figures
pressure air
the
installation
exit
The pressure
cooling than
for
drag. the
STD inlet.
The decrement
duced by the 0.6F pressure
pressure
conditions
parameters only
appears
with
the
in
exerts
air
in Figure engine
in baffle
at the
The effect
level
and
external
result
nacelle
the angle
in is
of attack
drop is
pressure
drop,
somewhat lower
in the pressure
influence
mass flow.
change.
dependency
same baffle
to reduce
on cooling
The 0.3F and GAC after
the modifica-
leakage. seen for
An increase the
The GAC inlet
0.3F in-
appears
to
of the 0.3A and 0.6F inlets.
of propeller 52.
pressure inlets,
baffles
of this
of altitude
51, were tested
pressure
shown in Figure
and 0.3A
the
for
50, the 0.3F and STD
an important
as cooling
as a result
and baffle
In Figure
recovery
These
accompanies
approximately
wbich
to the
function
which
The marked distinction
results,
of 0.2q_
the
loss
50 and 51.
This
in the
0.3A and 0.6F inlets
dra_ as well
is
on airspeed.
airspeed.
generate
recovery,
let
to its
shown in Figures
dependent
capability.
tion
are
at the exits
variation inlets
drop pro-
drop and plenum pressure
to be due to changes
pressure
inlet
due totally
are shown to be independent
slightly
while
is
pressure
recovery.
Baffle cruise
inlet
in baffle
operation
Improvements
on inlet in pressure
drop are obtained
while
STD shows a reduction
the GAC inlet with
performance
for
recovery
the 0.3F,
0.6F,
shows no change,
propeller
running.
and
The 49
behavior
of the
this
regard.
the
interaction
stream
STD and GAC inlets
It
is
believed
between
External Flow lets
visualization.
was
tained
for
a wide
range
propeller
the of
that
the
test
conditions each
The
The
propeller the
positive nose
angle
cowl
below
the
indicated is
shown
an
upward
50
flow
as
well
figures,
component
to
as
point spinner. spinner
inlets. at
the
GAC
are 53
the
were
ob-
also
inlets.
After
Initially, flows,
and
manner,
is
the
A and
to
strong a flow
spinner.
in
also
the indicated at
a The
immediately
upward
flow
outward
into
the
The
inlet
cruise
stopped.
condition.
strong
and
Figures
with
operating
be
the
propeller
shown
cruise
appears
climb
presented
As
and
observing
systematic
running.
in-
tuft
run.
is
slip-
in-flight
air
running
The
about
coolin_
a
nacelle for
of
in
with
of the
representative
Figure
the
attack
the
in
studies in
the
and
behaved
inlet
of
use
were
propeller
of
flow
data
0.3A,
with
propeller
below
the
reduced
stagnation
aft
external
were
stopped other
Effects
conditions,
tuft
STD
Inlet
conditions
in-flight
53-57.
in
flight
external
conditions,
geometry.
0.6F,
operating
component
distribution
0.3F,
understood
answer lies
inlet
through
Pressure
the
swirl
The
investigated
photographs.
that
the
and the particular
is not
is flow
inlet flow
has is
separated small with
on the
radii
the
following
distribution
the
evident
already
increases
The external Figures
flow
propeller the
54 and 55 is well
be primarily lateral
in the
ordered.
and 57 show behavior
similar
The stagnation
below
point
flow
below
stalled
the
for
inlets
is
taken. 0.6F, Pressure
both
inlets.
pressure where The
spinner
aft
inlet
pressure
and
inlets
for
separation Figures
56
causes upward flow
of the
spinner,
area appears
distributions.
GAC
no obvious
flow
and a to be
about
the
and orderly.
for
data
the
to
STD configuration.
The external
results
0.3A,
to the
in
in
appears
with
and no flow
The intake
unseparated
Inlet locations
inlet.
Stopping
The flow
configurations
immediately
inlet
area.
direction
components,
The 0.3A and GAC inlet
separation.
0.3F and 0.6F inlets
longitudinal
or azimuthal
on the nose cowl
the
local
running.
to
peak,
The inboard
separated
about
tends
a suction
inlet.
the
propeller lip
and subsequent
with
of
be seen
with
produces
on the outboard
the propeller
As will
on the upper
This
gradient
separated
as a result
results,
point
inside.
adverse
is
contour,
used there.
stagnation
move toward
This
intake
of curvature
the pressure
stopped,
is
lower
the
the
STD
Figure distribution
cruise are
inlet
58
were
data
condition
presented not
shows
in
for
the were the
Figures
acquired.
0.3F, 59-62. For
51
the cruise running
condition,
under
load,
The pressure given
59.
distributions
indicates
angle
of attack.
outboard
that
is
intake
side
points
duct
With
on the
consistent
to stopping
the propeller,
the stagnation
appears
to be small.
is
evident
is
due,
in part,
trimming
The 0.6F inlet While velocity 52
this
inlet ratio
was designed the 0.3F
and
of the
which
causes
than
the
stagnation
inside air
which flow
is
that response
no appreciable
move-
The change in velocity
is
It angles
shutting
This
which
that
resulted
this from
down the right
60 shows similar
inlet,
show
behavior
is believed
to operate
here
distributions
stopped.
inlets.
in Figure
than
inboard
shows little
The side pressure
after
sur-
at a positive
ratio the
indicating
to side-slip
the aircraft
the
velocity
inlet
point.
of the
inlet
and lower
in cooling
when the propeller on all
the pressure
governor
move to the
The outboard
ment of
asymmetry
the upper
stopped,
the reduction
this.
are
to be due to blockage
inlet
accompanies
36,
between
the propeller
with
0.3F inlet
are operating
at a lower
inboard
and feathered.
the axisymmetric
by the propeller
to operate
outboard.
inlets
believed
stopped the
with
between the
the propeller
to Figure
The difference
inlets
inboard
for
Referring
The difference
for
propeller
are consistent
faces
this
and for
distributions
in Figure
design.
data were taken
engine.
behavior.
at a higher
the pressure
distributions
do not
reflect
indicating
this.
a lower
the internal
Figure
operating
total
be stalled
They are more peaky than
pressure
in both
intake
35 no longer
to interpret
geometry,
stagnation
pressure point
similar
behavior,
for
lower
pres-
the axisymmetric of
the inboard
contour.
case. both
with
The stagnation
indicates
a local
stall.
geometry
involved,
it
The outboard Again, is
of 0.8q
.
contour
the
axisymmetric
for
is noted
outboard
inlets
showing
distributions the
stopped
moves towards lateral
because
impossible
The motion
lip
pressure
point
in
attached
some difference
models
pressure
indicate
and the
The side
inlets.
internal are given
data
Unlike
from the axisymmetric
for
similar
the external
recovery
the propeller.
exhibit
inside
point.
of separation
on the upper
both
propeller
with
pressure
configurations,
differ
inlets,
The results
The internal
stopping
the
impossible
demonstrates
a more extensive
was obtained.
and an initial
with
is
of
beyond this
is no indication
the GAC inlet,
62.
of the
it
61 also
to
flow.
distribution
flow
inlet
the results
distributions
consistent
and there
the external
Figure
is
0.3F,
However,
showed this
so that
the previous
sure distribution
ratio.
Accordingly,
in Figure
As with
For
ducts,
the pressure
behavior.
surveys
apply.
The 0.3A inlet
velocity
the
distribution
of the
to explain
the
complex this 53
behavior
further
or additional
without
an appropriate
experimental
analytical
model
data.
EXIT INVESTIGATION Background The exits
of the cooling
system throttle.
Since
the quantity
of cooling
the pressure
at the exit
flow
pressure.
volume
rate
and exit
duct
the flow
and lower
the
flow
settings
Exit
54
to the
local
between
rate.
cannot
be generated
a spoiler,
wake immediately
flow
primarily
area will
between
by throttle
an exit
the
cooling
can have an adverse which
configuration,
for
external
Too large
function,
required
configurations
an exit
so that
cooling
is governed
problems
volume
essentially
pressure
equal
the climb
The mechanism most often is
is
which
exit
pumping mechanism for the
itself
An additional into
throughout,
adjust
the volume
flow,
subsonic
will
Too small
in mixing
and external
corporated
is
act as the
flow
pressure
area.
area may result
downstream.
the flow
The relationship
the exit
flow
installation
flight
cooling
must be in-
is a cooling condition, at high
sized
is
a cow]
to produce
where
velocity. flap,
which
a local
low
downstream. and locations
flow
power
by the low flight
utilized
effect
are also
well
standardized. surface
The predominant
of the cowl,
external
pressure
Dependin_
is
on either
on the upper
close
surface
the
configurations
into
where
two ducts,
is
have been located for
twin-
a low pressure
region
available
for
pumping.
A negative
aspect
of the
surface
location
is
depositing
of oil
and grime,
is
picked
which
the
up by the
may come in
cooling
contact Exit
The original Figure
63.
This
cowl beneath flap
is
left
little
PA-41P exit
configuration
leading
cooler.
to maintain
fire
location
be used.
varied:
exit
cowl was occupied
by the
integrity,
Three exit cowl
flap
with
dictated
The central landing
gear
the requirement that
system parameters deflection,
lower
installation
variations.
wall
area,
The existing
in combination
crew.
The cowl
configurative
This,
airframe
shown in
on the
edge of the wing.
which
and/or
is
system located
deployed.
upper
of the
the passengers Configurations
room for
and oil
on parts
Test
shown fully
the lower
with
is a split
the
area of
flow,
one
may be used
particularly
there
static.
related
duct
exits
of the cowl,
lower
The local
and other
or a single
Periodically,
on the
to free-stream
may be split
of the cowl,
on the bottom.
is
the nose.
gear placement
the exit
side
back from
generally
on landing
considerations,
engine
well
location
existing could
and cowl
be
flap 55
aspect
ratio.
flection flap
The relationship
and exit
hinge
in
area
relation
shows two examples flap
deflection
where exit
between
depends on the to the
of this
the
duct.
exit
design
consisted
of exit
test
flaps
were relocated
fairings
were then placed
the area to its
0.75,
that
to produce
aspect
so that by fifty
value
this.
The exit
Three cowl
flaps,
and 0.55,
were tested
the
exit
and then
with with
aspect
areas
at the
This,
however,
resulted
in different
cowl
deflection
shown in Figure
56
the
lon_
high
angle. 66.
aspect
flaps
line
ratio
The three
The range
and short
hinge
is
percent shown in of 1.5,
exit
areas
so that
flaps all
same settings. flap
deflection
flap
cowl
in deflection is
reduced
The cowl
the
short,
Restrictor
ratios
matrix.
produced
the
PA-
each of the
three
with
original
area variation
a 3 x 3 configuration
same exit
The
to a fifty
the
between
first
ratio.
duct which
by relocating
largest
the
percent.
were installed
angles
im-
were investigated
flap
in the
original
below 65.
which
area and cowl
area was increased
Figure
cowl
and the other
dictated
parameters
41P exit
decrease
one where
be used.
The exit
cowl
64
The restrictions
posed by the PA-41P installation arrangement
de-
of the
Figure
area,
constant.
flap
location
arrangement,
increases
area remains
exit
cowl
having
flaps
the
are
angles
shown in Figure
67.
All
pertinent
geometric
data
Exit The results given
68.
the
lower
to the exit
drop from
inlet
to exit
condition,
means an increase modated,
figure.
in lower
completely
With
the
also
decreases,
reduced
is
with
clear
because
increase
andcorrespondingl_ creasin_ design
exit
of leaky
increasing
the
However,
internal
velocity
cooling
in cooling
increased
flow
is not
that
plenum pressure recovery
internal
that
drag will
It
pressure
can be increased
can be compensated
is
velocity.
baffle
is
con-
however.
and the flow
flow.
seen in this
drop,
ratio
area to achieve cooling
is
pressure
The implication
baffles exit
drop can be accom-
the upper
68, however,
area.
plenum pressure
pressure
inlet
pro-
any specific
plenum pressure
flow
the
with
from Figure
for
in lower
baffle
cooling
inversely
the combined pressure
an increase
the higher
losses
is
plenum pressure
into
increased
equation
pressure
The change in lower
verted
inlet
in baffle
and accordingly,
The variation
of Bernoulli's
constant
a reduction
II.
are
Since
is
Table
Results
plenum pressure area.
in
area investigation
Application
portional
flight
Area Test
from the exit
in Figure
shows that
are presented
drop,
by in-
poor inlet by subsequently
the required
cooling.
be increased
through
57
increase
in momentum defect,
creased
through
deleterious
mixing
effects
and external
with
the external
Figure
69.
Test
from the cowl
Due to cooling
at a low speed cruise
The prime
difference
slipstream
velocity
cowl
flap.
pressure flap
three the
position, flaps.
flap, tioning under
of the
in the
closed
linkage load.
in baffle
This
position towards
58
that
cowl
increased
is
the
the
it
of
lower
plenum cowl flaps
In the
that
all
as
moments on the The posi-
some defection
position It
in the results
due to the medium and long
the open position
additional
same for
increased.
II.
climb.
as the cowl
be the
allowed
in Table
spread
than
be exDectedo
closed
was
drop with
the hinge
flaps the
in
was observed
also
in
effectiveness
the curves
area should
position,
study
rather
pressure
increased,
somewhat above the values accordingly,
in
the test,
of the
this
absence of the
to what would
flap
are presented
show a decrease
the exit During
study
condition
the
The spread opposite
flap
influences
The results
is
length
is
which
deflection.
closed
here
and an increase
are closed
and associated
Results
requirements,
performed
the
flow
downstream. Cowl Flap
The results
drag mav be in-
cowl
due to aerodynamic
is
exit
areas
believed,
at the closed flaps loads.
deflecting The
cowl
flap
airloads
vanished
towards
the open position.
correct
for
as the
The Table
the no airload
cowl
deflection that
flaps
angle
exit
area,
not
mechanism here. to generate which not
deflection
for
the
The interest exits
is
in
driven cooling,
i.e.,
bottom
and exits
the
expedient, to the for
lower
twin
pressure
rather
engine which
aircraft,
the engine. flow
This ignores
negotiate
First,
surface the
supposed
effect
is
cooling for
flow
flow
air
engines enters
exits
with
the
are more
to return
again
Second, particularly
are obvious
pressure
the fact,
an adverse
is
downstream This
the advantage
region
flap
surface
exhausting.
the baffle
in a low pressure
ultimately
upper
seem to offer
pumping to increase
cowl
Investigation
ideas.
there
appears
tested.
than requiring for
it
the controlling
where the cooling
surface
the wide
II,
flow.
the use of upper
top,
are
seen between
immediately
cooling
Location
by two basic
updraft
is
of the
region the
is
in Table
configurations
Exit
values
Considering
angle,
The deflection
as a pump for
apparent
69.
as listed
a low pressure
acts
II
no difference
in Figure
range,
were deflected
open condition.
In the open position, the three
flaps
regions
of additional
drop across
however,
that
on an aerodynamic pressure
of low
gradient
the body,
must
to reach 59
free
stream
effects,
static
this
separation
is
in
pressure. difficult
region
downstream
corresponding
increases
Nacelle
pressure
locatin_
the
upper
Doints
static
bols
in
pressure sented
figure
data in
pressures
at
exits,
longitudinal The
low
wing are tween
the point
pressure
section for
near
cruising
the
nacelle
side.
Due
region
generally
60
in
71
to
the
belts
point the
the
were
on
the
the
presence higher
with
shows
positioned
the the
72.
on the
than
sym-
the
external and
suction
side the
prestatic
at
upper
difference
fuselage,
to
the
results
Figure
inboard
velocities
The
The
the
the
for
position
In
A noticeable
of
70
surface,
edge.
on
pressure
nacelle
lower
coincided
pressures
and
pressure
leading
Prior
measurements.
71
show
flight.
has
and
Figure
indicated
72
lowest
nacelle
taken.
Figures
and
of
in severe
seDaration
exits,
were
relate
Figure
in-
measurements.
distribution
given
flow
in drag.
pressure
pressure
the
of flow
surface
measurements
longitudinal
result
distribution
distribution where
flow
of circumstances.
can well
in terms
layer
without
of a low momentum secondary
to a low pressure
to
to achieve
the most favorable
The introduction
effects
Due to boundary
the
surface. peak
71,
of
the exists
and flow
comparable
the in
the
results beoutboard this points
outboard
of the nacelle.
aircraft,
the
angle
swirl
of attack
the angle
results
climb
cluded
for
peller
stopped.
I00 kts,
dicated
by the
cruise,
there
sure
the
at
Exit surface the
exit the final test
stream
exits
location is
largest
of
investigation exits test
to
align
the
climb
surface.
the
bv
results. shown
in
the
three
(150% the
lower
them
Figure
exit
the pressure pressure
is exits
is
on
the
increases
the wing
as in-
73.
The
shown were
in
of
the
area
the
Louvers
was
were
external
Figure
74. off
upper
lower
the
closed
nres-
surface.
exit in
original). with
In
upper
installation
by
found
of the suction
investigated
flow
inpro-
top of the nacelle.
The
of
for
The pressure
doubling
locating
the
speed of approximately
lowest
the
also
lower
Also
same data
cases,
a potential
This
72 presents
as one moves away from
configuration
program,
are the
the
and decreases
airspeeds.
amplifies
for
increases
section.
Figure
slipstream
wing
results is
exits
section.
At the best
to the upper free
wing
of the
and correspondingly
purposes
In both
towards
section,
power at different
comparison
0.4 q .
wing
velocities
side
slipstream
of the outboard
the propeller
adjacent
at
inboard
on the inboard for
about
of the
to higher
pressures
on the starboard
of the propeller
of attack
contributes
Also,
and
set
surface used
flow. During all
at The
the cooling
61
flow
passed through
in Figure sets on,
of data louvers
upper the
off,
engine
exits .
condition,
However,
by deployment
of the
are shown to be superior
climb
to the conventional
in climb
expect than
the
A drag study surface
exits
of climb results
the upper lower
was performed
in both
were inconclusive. the relative
eralized
speed/power
uration.
The results
generalized
aircraft.
62
climb
exits.
system with comparing
for
in a definite
In terms
of generalized
are
have been surface in
Intuitively,
system to offer cowl
less flaps
drag open.
the upper
and lower
configurations.
Rate
the
climb
test.
could
be rendered
the two systems.
in Figure
generalized
result
exits
In
and comparable
method was used for are given
drop.
The upper
No judgement drag of
power versus
exits
flaps.
resis-
systems
lower
and cruise
was used as the measure
concerning
surface
surface
some flow
exit
surface
surface
The
drop across
pressure
in cruise
lower
louvers
exits.
pressure
and lower
cowl
with
surface
baffle
the
Three
exits
contribute
here
exits
one would
surface lower
obtainable
are given
conditions.
the baffle
the upper
equal.
modified
upper
The louvers the
The results
and climb
increase
affects
essentially
cruise
exits.
and conventional
by 0.1q
which
climb
both
are presented:
surface
tance the
75 for
the upper
The gen-
the cruise
config-
76 in terms
velocity.
of
The upper
drag increase velocity,
The
there
for is
the
approximately setting.
a 6 kt
During
the horizontal exits.
With
noticeably, Tuft
the test tail
the
louvers
Other
the
formation
mined.
However,
support
to the
uration
tested, surface
crease. wind
ment with dicates
exits
lead
is
included
indicated
in the
results
to date,
cooling
air
exits
support
in low pressure
wake.
buffet, deter-
of the wake lends For
the
benefits
configof
the
drag in-
in a full
The drag results
evaluation.
tail
by an associated
scale
were in agree-
Reference
in a low pressure
configuration
do not
developed
change in flow
was tested
the exits
surface
wake was not
results.
here.
in
increased
through
installation
reported
was felt
the upper
of a well
presence
II).
same power
buffet
the buffet
of this
configuration
to the best
the
any significant
were negated
locating
from
manifestation
cooling
the results
not
its
(reference
that
removed,
and nature
the
a mild
flow
drag measurement
A similar
tunnel
program
indicate
than
the
speed for
the presence
did not
patterns. actual
in
due to the
indicating
studies
upper
decrease
1 also
in-
region
when the resulting
does
drag
In summary, the reported the
apparent
regions
benefits
of
of the aircraft°
63
CONCLUSIONS
I.
With
present
cooling
tunnel
test
Much of the
to horizontally-opposed
installation aerodynamic
total
pressure
the
oncoming
flow
the
component passing
Current
design
for
the
external
leakage,
cooling
A simple
ground
an important cooling A flight installed
ground
through
practices
in regard
rather
than
for
the
by
static
should than
a rubber
results
in
and increased
blower
for
be of that
of
the engine.
baffle
problems,
to
are accounted
rather
of using
engine
tool
configurations
The measurements inlets
is
engines.
cell
measurements,
from the
test
test
measurements
engine
correlation
configurations,
measurements.
through
concerning
applicable
pressure
64
technology
directly
using
6.
engine
between
flight
practical.
and altitude
The differences
of
be performed
characteristics
internal
5.
through
orifice
and flight
4.
should
whenever
radial
measurements
to obtain
Such investigations
wind
3.
reliable
drag are difficult
test.
2.
techniques,
tape
lap
seal
significant cooling
drag.
system has been shown to be development
of aircraft
installations. test
technique
engine
cooling
for
the
determination
requirements
of the
are determined
in
terms
of
airframe ground .
,
The
easily
measurable
manufacturer test
cell
aerodynamic
in
the
is
an
The agree
design
study
data
with
of
need
theory
region the
in
of for
The should
of in
of
consequences,
the
inlets
basic
inlet
locating be
freeing
the
imposed
by
regard.
cooling
exit
thus
restrictions
the
regard
not
the
this
behavior
parameter
installation. sure
from
effectiveness obvious
parameters,
area
are
factor
installation. design has
its
effect
of
the
exits
There
guidance.
been
to
attempted
a major
shown on in
without
the
to cooling
a
low
pres-
a
thorough k
65
TABLE I.
- COOLING DRAG FLIGHT TEST RESULTS
Configuration No cooling Augmentor Augmentor Augmentor
(+4%) (+0%) (+7%)
- COWLFLAP TEST CONFIGURATIONDATA
Cowl Flap Position
Exit Area (cm2 )
1-closed 2 3 4
214 268 300 326 366
5-open
0.0226 0.0234 0.0226 0.0242
flow 25% open 50% open 100% open
TABLE II.
66
CD @ CL = 0
Cowl Flap Deflection Short
Medium
13 22 31 42
5 I0 14 18
(deg.) Long
3 6 9 13
REFERENCES
I.
Hammen, to
T.
F.,
Installation
Engines. March 2.
H.
and
Tests
Cooling
Nichols,
M.
the
Nielsen,
V-770-8
Vol.
Aircraft
54,
No.
3,
MR
R.;
R.;
the
Bell
M.
Engine
Model
33
Airplane
A.,
Jr.:
An
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Installation NACA
Dennard, Engine
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WR
Keith,
the
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for
L-561,
S.:
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the
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1945.
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L5112b).
and of
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L5112).
and
MR
Cowling
J.
A.,
ll-Aerodynamics.
NACA M.
H.
1944).
1-Cooling.
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High-Altitude in
Pertaining
1945.
Emmons,
Propeller-Research 6.
Factors
Aircooled
Wilson,
a Fleetwings
May
V-770-8
(Formerly
of
In-Line
and
L-632,
and
NACA
Ranger
Nichols,
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(Formerly
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(Transactions),
of
MR
N.;
Ranger
XOSE-I
WR
NACA R.
the
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NACA
Conway,
the
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1946.
(Formerly
4.
Rowley,
Inverted,
Journal
Ellerbrock,
of
and
of
SAE
Flight.
3.
Jr.;
NACA
Schumacher, of
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the
Jr.: Airplane MR
L. Ranger NACA
An
Investigation in
November E.:
1943.
Analysis
SGV-770 CMR
the
October
of
D-4
the
Engine
1943.
67
7.
Monts,
F.:
The Development
Installation
8.
Data for April
Miley,
Cross E. J.,
Lawrence,
D. L.:
and Cooling
9.
1973. Owens, J. K.;
An Investigation
and
of the Aerodynamics
of a Horizontally-Opposed
Engine
Installation. 1978.
Miley,
Cross,
Jr.;
Owens, J.
and
S. J.;
D. L.: Engine
August
1977.
Miley,
S. J.;
E. J.,
Aerodynamics
Cross,
Corsiglia,
AIAA Paper 77-1249,
E. J.,Jr.;
Ghomi, N. A.;
Determination
a Horizontally-Opposed
of Cooling
Aircraft
Transactions,
Vol.
Katz,
J.;
Scale Wind Tunnel
Study
of Nacelle
Rubert,
August
K. F.; andKnopf,
of Cooling
Systems for
NACA WRL-491,
Mass Flow
Installation. 1980.
R. A.:
Full-
Shape on Cooling
Drag.
1979.
G. S.: Aircraft
A Method for Power-Plant
the Design Installations.
1942.
Biermann,
A. E.:
Engines.
SAE Journal
No. 3, 1937.
Engine
and Kroeger,
and
Air
88, August
V. R.;
AIAA Paper 79-1820,
K.;
of Horizontally-Opposed
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P. D.:
SAE Quarterly
68
Jr.;
86, September
for
13.
SAE
Vol.
Bridges,
12.
Aircraft.
Transactions,
Aircraft
Ii.
Aviation
Engine
SAE Quarterly
Lawrence,
I0.
General
Paper 730325, S. J.;
of Reciprocation
The Design
of Metal
(Transactions),
Fins
for
Vol.
41,
Air-Cooled
14.
Goldstein, bility
A. W°; and Ellerbrock,
and Heating
of a Baffled 15.
Neustein,
Across
Pinkel,
B.;
and Ellerbrock,
Cooling
Data
E. J.:
Valerino,
High-Altitude
Flight
Cooling
Engine.
Kuchemann, D.;
Stockman, for
Calculation
Inlets.
and Weber,
N. O.;
M. F.;
Inc.,
and Button Potential
NASA TM X-68278,
Correlation
1940.
and Bell,
B. B.:
of a Radial 1947.
Aerodynamics
S. L.:
of
and Several
No. 683,
No. 873,
New York,
of
Loss of
Investigation
J.:
1944.
Cylinder.
Cylinder
NACA Report
Book Co.,
Comparison
Jr.:
NACA Report
Manganiello,
No. 783,
the Pressure
from an Air-Cooled Engines.
Jr.:
H° H.,
Compressi-
Loss and Cooling
Aircraft-Engine
No. 858, 1946.
McGraw-Hill 19.
a Baffled
Jr.:
NACA Report
L. J.,
NACA Report
Air-Cooled 18.
Barrel.
and Schafer,
Multi-cylinder 17.
on Pressure
Methods of Predicting
Altitude
16.
Cylinder
J.;
Several
Effects
H. H.,
of Propulsion.
1953. Computer
Flow in Propulsion
Programs System
1973.
69
I
J
(a)
(b)
Figure 7O
I.
- Aircraft
radial
in-line
engine
cooling
installations.
"--. ', ©...o..o...O..O...q
(c)
vee
I
(d)
horizontally-opposed
Figure
I.
- Concluded. 71
C) Q; Q;
f_
bO
r..4 0 0 C;
.r.4
Q;
4--} q-4
(.;
pC_ I
-,.n'o")
o Q; pc_
c,,l Q;
b_ .,.4
72
4J _4 CJ
4J
!
!
b_
73
5
Figure
74
4.
- PA-41P
test
aircraft
cooling
installation.
co I-
_7 _I:I,I I,III: -r_ hl n"
o o
I ILl
r,-_
o
o
i-i l.iJ _: i:I.
t_l\ c'xl\
tl:)_ e_l\
Ill
o
o _
,,=
i i\
\i_
i,_
-I:I. 0
I-n_
'
Z
,.o
oo oas / 8_l - MO-I:I _IIV 9NI-IO00 0
._i
5
,,7
Io
0 0
o
\ 0
0 0
o
0
_ 4-1
i_
0 0
o 0
0 o
_
bO
I
._o.N ,H
"'°
_
.0
75
i
®
®
----
®-___
----
____
___
®
---®
Figure
Figure
76
7.
6.
- Cooling
installation
model
schematic.
- Engine orifice characteristics and cooling requirements determination test set-up,
KIEL
TUBES
I
!I
I
I i]Jilll
_OTOT_ilitl IIIIIIII I I D D ifI,,lllllllJ
I ENGINE TOP
VI EW
lllli_JillZ I I n
lllt;t_
0
_,lJllll® lllltIIIliil D I I I
I
I
- Location of temperature
I I iiIIillllll_
I n n ,,II JJitII IIIIIi ®
PICCOLO TUBE
8.
ii
llilll_
iilll,,iiiiii®ll-ill
Figure
FI I
Kiel total probes in
TEMPERATURE THERMOCOUPLE
pressure the high
probes pressure
and plenum.
77
4--) ,--M "H
4-J
.r-I
O 0
t_
O
(D
g} (9 _L
4-J O 4.J
(D
I
,r--I
78
bO .,-I
4-1
0
4-3
4J "l:J
UI
12u r-4 4-1 0 4-1
11
4-1
01--t 4-1
4-1
m
_-4
12U
I
o
(1)
o_
79
,,.,,IILIIIIL
[illl ' II*I
ii
(a)
Figure
80
II.
probe
.
I iii_
,,.,,fILl Will
locations
- Internal flow pressure probes, locations, cylinder numbers. Circled numbers denote type; dots denote locations.
and probe
9.5 mm
(b)
probe
vertical
positions
INTERCYLINDER BAFFLE
"_
..........
T_ BAFFLE
oo.°°
.....
Dt,°,°°,,°°=,,_
UPPER LOWER PLENUM PLENUM
/
BU
I
PROBE (_'_1.
'
.lrT BAFFLE-SI41ELD-UPPRoBE BAFFLE
(c)
baffle
button
Figure
and
II.
lower
plenum
-SHIELD-DOWN PROBE
static
probes
- Continued. 81
Scm
DRILL
#
-_
60 HOLES
5cm
7
ON
5 cm
SPAClN_'_'_G
(d)
...
(e)
,
static
r-
L. ¸
,
piccolo
tube
.J
pressure belt location r____
FIN-SHIELD
detail
......
/
PROBE _.1
7
(f)
___t
fin-shield
9.5ram
A I
Figure 82
II.
- Concluded.
probe
installation
C) BAFFLE-SHIELD-UP PROBE [] BAFFLE-SHIELD-DOWN PROBE FIN-SHIELD PROBE
I
I
1.5
I
I
I
20 PICCOLO
Figure
12,
- Comparison of measurements.
I
I
2.5
I
3.0
I
I
I
3.5
kN / m 2
different
lower
plenum
pressure
83
¢¢3 O
t I
0
0
bl
O0
D
0
g
0
o O
OI
O I
0 --:
0
[]
o:
i C,I
0
UO_O
0
0
0
1 I
--
w
I
C
0
0
0
0
0
d
ol
t
LLJ nn :E
©
k_
n_
01 01 0
Z 1 )0 ,@-
0
P
5 ._ I
I
n_ 0
n,w I
0 J
iO0
ilO
EQUIVALENT
Figure
142
68.
-
Effect of and engine
120
130
AIRSPEED
140
150
- KTS
exit area on lower plenum baffle pressure drop.
pressure
O
& 0 n" a
0.7
LLJ nOr) O0 O.6 LLJ nO_ ILl d LL LL. 0.5 m
COWL FLAP
a. 0.4 0
0
SHORT
[7
MEDIUM
/:', LONG
!
W nZ_ Or) CO 0.3 W n-
:E Z 0.2 LU J {3. n" L[J 0 _J
0.I
I
I
I
I
I
I
2
3
4
5
CLOSED
OPEN
COWL FLAP POSITION
Figure
69.
-
Effect pressure speed
of
cowl and
cruise
flap engine power
aspect
ratio
baffle
pressure
flight
on
lower drop
plenum for
low
condition.
143
,--I ,--I (D
.I-J
0
_0
r_0 0 .,-I IJ
0 i.--I
,--I ,.Q
v
Ia-,
.,--I 4.1
_1 _9
°l
0
°,-I 0
0
144
NACELLE 0 4) []
I_. -1.6 O I
ILl A" :D O3 03 -1.2. LLI n," 13. (..)
POSITIONS
Inboard side Outboard side Lower (average)
Upper
(?y..
-0.8 O3 W -J -J W 0 ra
tt" O3 O3 U.J n," Q-
for
"0 -I.2
_.
Inboard side Outboard side Lower (overage)
-
_-.
UpperCL. _
I'-o3 w J J w
-0.8
-
z -0.4
[] .....
Figure
72.
-
'
I
I
I00
liO
120
0--
i
130
i
I
140
150
EQUIVALENT
AIRSPEED
- KTS
Nacelle
static
pressure
for
climb
power.
145
.4-) .r-(
0..) r,.,)
I'.-4 CY')
(::)..,
0
F_ O .1_ i,-I l--I c_ .I-I c_ I-I
c_
/
146
(a)
inboard
(b)
outboard
side
side
Figure
74.
- Upper
surface
exits. 147
O
[]
!
n 0.7 0 n," 0 ILl n" 0') 0.6 CO W n,n W ._J h 0.5U