An Experimental Investigation of the Aerodynamics ... - Contrails

Turbo. West. Corporate. Aircraft. Center. Broomfield,. Colorado. Prepared for ...... turbo- supercharging, the same four power settings were obtainable at all ...... b.I 0.09. O/.
8MB taille 2 téléchargements 263 vues
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

Investigation

Installation NACA

Dennard, Engine

J.

WR

Keith,

the

and

Cooling XP-77

for

L-561,

S.:

An

Installation

Bell

A.

in

the

Edo

1945.

Investigation for

NACA

L.,

XP-77

Tunnel.

N.;

Cowling

the

WR

of

Edo

L-562,

1945.

L5112b).

and of

Jr.:

L5112).

and

MR

Cowling

J.

A.,

ll-Aerodynamics.

NACA M.

H.

1944).

1-Cooling.

Airplane.

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,

Jr.;

Airplane.

(Formerly

5.

H.:

(Transactions),

of

MR

N.;

Ranger

XOSE-I

WR

NACA R.

the

XOSE-I

H.,

NACA

Conway,

the

W.

1946.

(Formerly

4.

Rowley,

Inverted,

Journal

Ellerbrock,

of

and

of

SAE

Flight.

3.

Jr.;

NACA

Schumacher, of

Airplane.

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

Installations.

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