Lockheed Field Service Digest FSD Vol.3 No.6 Intro L1649 Starliner Part 2 of 3

8/18/2019 Lockheed Field Service Digest FSD Vol.3 No.6 Intro L1649 Starliner Part 2 of 3

1/32

MA

JUN

9 7

V

N


2/32

May June 1957 Vol. 3

No 6

COVER PICTURE: M o re f r ei gh t

a t

a lo we r ra te .

In

pointing

toward

this

goal,

Th e

Flying

Tiger line,

world s largest

freight

and contract carrier, ha s or

dered

10 of these

ne w Model

1049H Super Constellations.

Available

in two distinct config-

urations,

the

Super shown

here

is

completely flexible for

cargolp as se ng er o pe ra tio ns

an d

is

convertible from cargo

to

mixed, medium,

or

high den

sity

passenger

service.

The 1049H f reighter config-

uration not shown

is

the largest

an d fastest all-cargo airplane

in

the world, carrying

payloads in

excess of

21

tons at

335

mph.

Terence B. Donahue Editor

ONT NTS

T H E STARLINER PART

II

POWER

P LANT

FUSE L AGE .

EMPENNAGE

G R OU N D H AN DL IN G PROVISIONS

NOSE

GEAR D O W N

LOCK

MECHANISM

SCOVILL

ADJUSTABLE FASTENERS

.

D - C GENERATOR OVERVOLTAGE

RELAY

PANEL

LANDING

LIMITATIONS

W I T H T IP

TANKS

INSTALLATION OF LEATHER BACK-UP RING IN BRAKE SELECTOR VALVE

NOSE

GEAR STEERING RESPONSE .

PROTECTING INTEGRAL FUEL

TANKS FROM

CORROSION

OIL

FOAMING

IN AIRESEARCH

CABIN

SUPERCHARGERS

W I N G F L AP S HO RT LINK ROLLERS

TRADE TIPS

INSTALL O IL DEFLECTOR AT

ENGINE

OVERHAUL

EASIER

ACCESS

TO LUBE

FITTING

COMMERCIAL

SERVICE BULLETINS

PENDING

TECHNICAL PUBLICATIONS FOR TRANSPORT AIRCRAFT

LOCKHEED

A IR CR AF T C OR PO RA TIO N

3

6

8

8

12

14

15

15

16

19

24

25

26

27

27

8

The

Lockheed Field

Service Digesl is

p ub li sh ed b im on lh ly b y

Lockheed Aircraft

Corporation, Califarnia

Division, Burbank,

Califarnia.

No

material

is officially approved

by the

CAA, CAB or a ny

of the

military

services unless specifically

noted.

Airline an d military

personnel

ar e

a d vi se d t ha t d ir ec t

use of th e

informa

tion

in

this

p u bl i ca t io n ma y

be restricted

by

directives in

their

arganizations. Obtain

written

permission

f rom Lockheed Aircraf t

Corporation

before republishing

an y

of the

material

contained

herein.

This

require

ment is

m a nd a to ry t o e n su r e

Ihat all

material republished

will conform to

t h e l a te s t

information an d

changes.

The fol lowing marks

ar e registered

an d

o wn ed by

Lockheed Aircraft

Corporation: Lockheed

Constellolion,

Slarliner lectra Lodestar and

peedpak

Address

all communications to Lockheed Aircraf t

Corporation,

Burbank,

California;

Attention, Field Service an d Training Division.

COPYRIGHT

95 7 BY

LOCKHEED AIRCRAFT

CORPORATION BURBANK CALIFORNIA


3/32

TH

P RT

W

T

HIS is the second in a series of articles describing

the features

of

the new Starliner. In an article

which appeared in the preceding issue of the igest

Vol.

No 5

we gave a general description of the

airplane pointed out the service areas and described

the new wing.

In this issue we will describe the power plant

fuselage empennage and ground handling provi-

sions for the new airplane. In a future issue we will

discuss the landing gear flight controls and func-

tional systems to conclude our introductory presenta-

tion of the Starliner. However items of interest

concerning this new airplane will be reported in

future issues

of

the igestconcurrently with news

of

Constellation series aircraft.


4/32

SPE IFI TIONS

PERFORM N E

D T

MANUFACTURER WRIGHT

AERONAUTICAL DIVISION GUARANTEED POWER

RATI

NGS

ALTITUDE MAP

MODEL

988-TC18EA-2

TYPE

18

CYLINDER, DOUBLE ROW,

TAKE-OFF

BHP

RPM

FEET IN

HG SFC

AIR COOLED

RADIAL

SEA LEVEL 3400 2900

SEA LEVEL

58.5 0.68

BORE

6.125 IN.

LOW

RATIO 3400 2900

4 000 56.0

0.674

STROKE

6.312

HIGH RATIO 2550 2600

15 200 49.0 0.632

TOTAL

DISPLACEMENT

3350

CUBIC

IN.

MAXIMUM CONTINUOUS METO

COMPRESSION RATIO

6.70:1

SEA

LEVEL

2800 2600

SEA

LEVEL

51.0 0.654

SUPERCHARGER RATIOS

6.46:1 AND 8.67:1

LOW

RATIO

2850 2600 4 700

49.0 0.645

IMPELLER DIAMETER

13.5

IN.

HIGH

RATIO

2450 2600 16 400

47.0 0.628

PROPELLER REDUCTION

GEAR

RATIO

0.355:1

ALTERNATE MAXIMUM CONTINUOUS

PROPELLER

SHAFT ROTATION

CLOCKWISE

SEA

LEVEL

2860 2650

SEA LEVEL 51.0

0.658

CRANKSHAFT ROTATION

CLOCKWISE

LOW

RATIO

2920 2650 4 800

49.5 0.650

POWER

RECOVERY TURBINES

3

MAXIMUM RECOMMENDED

OVERALL LENGTH

89.53

IN.

OVERALL DIAMETER

56.59

IN.

CRU ISE POWER

MASTER

ROD LOCATIONS

NO. 1 AND

NO.2

CYLINDERS

LOW

RATIO

1910 2400

13,500

33.5

OIL

PRESSURE

70

±5 )

PSI

HIGH RATIO

1800 2400

22,100

34.5

FUEL

SUPPLY PRESSURE

24

TO

26 PSI

CRITICAL

ALTITUDE AT

BEST POWER MIXTURE

figure 1

The ·2

ngine


5/32

POWER PL NT

Each power

plant

consists

of

an engine complete

with propeller, accessories, related systems, cowling,

and mounting structure.

The

power plant joins the

nacelle at the firewall where disconnects are pro

vided

for

electrical, plumbing,

and

control systems.

By transferring certain accessories and

by

making

minor changes in some

of

the plumbing, a power

plant may be installed in any of the four power plant

positions.

ENGINE The 1649A is powered

by

the Wright Aero

nautical Model 988TC18EA-2 18 cylinder, radial,

air-cooled, turbocompound engine. A single stage,

two speed, gear-driven supercharger is contained in

the engine aft section. The two-speed supercharger

control is actuated by a new adjustable cable system

with a built-in spring capsule shock absorber.

The EA-2 engine incorporates many durabi li ty

improvements over the DA series in all sections of

the engine.

The main engine oil tanks, located just

af t

of each

firewall, feature

Winslow

full-flow scavenge oil

filters. The filters can be serviced through access doors

in the upper

part

of the nacelles.

The planetary reduction gear system in the nose

section of the engin e provides a propeller-to

crankshaft ratio

of

0.355 to 1 compared to 0.4375 to 1

on the 1049G. Thus the prop-tip rotational speed is

reduced by approximately 10 per cent from that

of

1049G propellers. Yet the large diameter propellers

on the Starliner provide greater thrust. This combina

tion

of

a large diameter propeller operating

at

relatively low rpm results in more efficient and quieter

propeller operation. Figure 1 shows the EA-2 engine

and lists specifications and performance data.

PROPElLER Either Hamilton Standard Hydromatic,

or Curtiss-Wright electric propellers can be installed

on theStarliner. The propeller control system provides

the fo llowing features for both of these propellers:

constant speed governing, synchronizing, individual

selective increase or decrease rpm, manual and auto

matic feathering, and reversing. Synchrophasing is

also available as a part of the Hamilton Standard

propeller control system.

ydromatic propeller installat ion Model 43H60-349

synchrophasing or

Model

43H60-355 synchroniz

ing are the Hamilton Standard propellers available

for the 1649A. Both are 16 feet 1 inches in diameter

and have three hollow

or

solid

Dural

blades.

Dual

feathering lines, connected to a spring-loaded selector

valve and incorporating spring-loaded shut-off check

valves, can be provided to furnish an alternate

feathering line in case of failure of the normal line.

urtiss electric propeller Curtiss-Wright propeller

assembly Model C634S-C602 is also available for the

1649A. It is the same diameter, includes the same

basic features as the

Hamilton

propel ler, and has

three extruded hollow steel blades.

POWER PL NT

MOUNTIN

The

engine mount

is

attached by four internal wrenching bolts and two

shear fittings to the nacelle structure

at

power

plant

stat ion 0.0. Four hex head bolts through the nacelle

ring, shroud, and oil cooler

af t

structure complete

the power plant to nacelle attachment. Six Lord Dyna

focal suspension mounts support the engine in the

mount ring. Although the mount is identical in

geometry to that used on the 1049G, the wall thick

ness of upper and lower support tubing and the

diameter of attachment bolts have been increased to

withstand higher loads.

ENGINE EXH UST SYSTEM The location of the power

recovery turbine PRT exhaust outlets is approxi

mately 30 inches farther forward of the wing leading

edge than on the 1049G. All PRT flight hoods have

been redesigned to allow their exhaust gases to pass

above the upper surface of the wing.

The

outlets for

the No.1 and No.2 PRT tail pipes are approximately

17 inches above the power

plant

thrust line. A steel

support attached between the engine and the No. 2

PRT tail pipe protects the PRT nozzle box from

undue loading caused by tail pipe deflections. Turbine

hood clamps are

not

changed from those used on the

1049G.

ENGINE OWLING Refer to Figure 2 The primary

engine cowling is

comprised of these major assem

blies: the upper panel, two side panels, the lower

panel assembly, and a removable oil cooler air scoop.

Flame shields are provided at the turbine exhausts.

Continued

1

page 6

figure

2 ngine owling

3


6/32

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8/32

Except for the changes necessitated by the new

location of the PRT exhaust outlets, the cowling

on

the 1649A is very similar to that used on the 1049G.

For instance, the action of the panels, cowl pane l

latches,

and

support rods remain the same.

Upper and

lower cowl panels are practically identical to the

1049G cowl panels.

The

left side panel is unchanged,

but the right side panel has been altered to accom

modate the changes in the

No.1

PRT

hood

and tail

pipe. Cowl flaps are of new design. Cowl flap actu

ators are similar to those used on 1049G airplanes

except that the stroke is slightly longer and end

fittings have been revised due to turbine hood changes.

EMERGEN Y

EXIT

RELE SE

. . . . . . TO

OPEN

Figure

ype

mergency

xit

oor

INTERIOR

RElE SE HANDLE

Pulls Up

EXTERIOR

RElE SE

PUSH PL TE

6

FUS L G

The fuselage is the same length

as

the Model

1049G, and is structurally designed to permit a pres

sure differential allowing 8,000-foot cabin altitude

at 25,000-foot airplane altitude.

The station number designations F 0.0 to F 128.8,

527.6 to 658.4, and A 0.0 to A 92, which applied to

1049C through G airplanes, are

not

used for the

1649A. Fuselage station numbers in the new fuselage

midsection are preceded with the letter M and run

from M 0.0 to M 351.6 see Figure 3 .

STRUCTUR L

CH NGES

As a result of the higher

gross takeoff weight of 156,000 pounds, the majority

of skin, stringer, and frame gauges have been

increased. All fuselage skins are of 2024-T3

or

2024-T4 clad material. Changes have been made in

the forward fuselage section, particularly in the keel

sons and in the FS 205 attachment fittings and bulk

head ring the ring is now a chemically-milled part .

Minor structural changes have been made in the

remainder of the nose wheel well area to withstand

the increased shear loads.

Mid-fuselage barrel section 4 F.S. M142 to F.S.

M247 has been redesigned to accommodate the new

wing, and is now completely cylindrical in cross

sec-

tion. Barrel section 5A, immediately

af t of

section 4

has been redesigned in order to fair in with the new

contours

of

section 4.

Unpressurized service areas are provided forward

and af t of the wing. See the illustration 1649A

Maintenance and Service Areas in Vol. 3

No.

5 of

the igest

Because of the extensive clearance between pro

pel le r tip and fuselage, it is unlikely tha t fuselage

skin damage will occur from prope ller ice. There

fore, the ice shields previously used on Constellation

models are not installed on the Starliner.

C IN DOORS N WIN OWS

The passenger door

and the crew exit door on the

left

side

at

F.S. 296

have been revised to include many improvements in

operation and appearance.

Cabin windows are Sierracin 611, the same type as

used in the 1049G.

EMERGENCY

EXIT

PROVISIONS

A Type I emergency

exit door is provided at FS 897 on the right side of the

af t

fuselage see Figure 4 . The opening is 24

by

48

inches and the lower sill

is

at

cabin floor level. The

door is held in position by a latch at the top center

of

the door and two fixed pins at the bottom. A handle

which

pulls up

releases the latch from the inside and

a push plate releases it f rom the outside. The door

opens inward.


9/32

Four type

V

emergency exit windows which open

inward

are

installed in the cabin window plane two

on each side of the fuselage directly over the wing.

he

latch mechanisms

on

these exits are the same

type as that used in the Type I exit door except that

the inside release handles pull down

as

shown in

Figure

he

cut outs for the Type

V

exits are reinforced

with forged aluminum alloy corner pieces incorporat-

ing a 4 inch radius.

LOWER

BAGGAGE

COMPARTMENTS

he

forward

and

af t

baggage compartments are similar to those

of the 1049G.

he

volume of the forward compart-

ment

is

239 cubic feet.

he

volume

of

the

af t

com-

partment is 317 cubic feet if one loading door is used

anfl 280 cubic feet

if

two doors are used. Zippered

compartment l inings are used in the same fashion as

in the 1049G.

PRESSURE

SEAL

BLANKET SEPTUM) o

form a pres-

sure seal and to eliminate the possibility of fuel vapors

reaching the cabin in the event

of

a fuel leak in the

wing center section tank a vapor proof seal or

septum is installed between the upper surface

of

the

wing

and the fuselage. This septum is a pliable

blanket

of

two ply nylon neoprene construction.

t

is

a ttached to the fuselage under floor structure and

covers the area shown in Figure 6. Spanwise ribs built

into the septum permit a venti lating air flow between

it and the wing upper surface. A sealed zipper runs

chordwise

at

the center

of

the two piece septum mak-

ing it possible to quickly check thewing upper surface

and the

wing

joint.

he

two halves

of

the septum can

be detached and replaced with relative ease.

ontinuedon next page

Figure

5

Type

IV

Emergency

Exit Window

EMERGEN Y EXIT

RElE SE

TO

OPEN

EXIT RELEASE LATCH INTERIOR RELEASE

HANDLE Pulls Down)

EXTERIOR

RELEASE PUSH PLATE

WINDOW

CENTERING PIN

TANK AREA

I

FORMER

TO PREVENT

I

I

F

S

FORW R

F S BALLOONING OF

SEPTUM

FT

DURING

AUXILIARY

AIR

PRESSURE

FORCING SEPTUM F S F

M 9 Z 0 S ~ R R ~ ~ C E ~ 1 4 4 3

VENTILATION

AGAINSTTANK

SURFACE

DURING

Ml44.45-SEARRVE ACE--Mlis

I

I

NORMAL PRESSURIZATION I I

SEPTUM

\ ~ ~ r i U i ~ ~ m m m m

FROM - _

VIEW LOOKING INBOARD

. . . . . . . . ]

AIR INLET AIR OUTlET : :

AIR OUTLET SEPTUM

FORW R FWD

AIR

OUTLET

SEPTUM FT

DRAIN AND VENT DRAIN

N

VENT FROM

AIR INLET

Figure

6

Pressure

Seal Blanket Septum

7


10/32

EMPENN GE

The

empennage does

not

differ appreciably

from

that of

the 1049G.

The

major change was made in

the horizontal stabilizer center section to accommo

date the new rudder and elevator booster units. This

involved changing the booster attachments

and

back-up structure installing new ribs

and

making

other minor s tructural revisions which were neces

sitated

by

the increased booster loads. Front

and

rear

horizontal stabilizer beam webs have been strength

ened

by

adding doublers in some areas and

by

increas

ing the web gauges in other areas.

Minor

changes in

the horizontal stabi lizer include gauge increases in

some corrugations

gauge

increases in a ll nose

hat

section stiffeners and larger rivet diameters.

The

torque tubes and supports

for

center and out

board rudders have been strengthened to withs tand

higher loads.

All rudder counterweights have been moved from

the control horns to the top parts

of

the rudders below

the top hinges. Cutouts have been made in the fin

structure to clear the counterweights.

Because

of

the increased empennage loads on the

1649A the four bolts attaching the stabilizer assem

bly to the fuselage are

of

higher heat treat material

than the bolts used

on

previous models.

GROUN H N LING

PROVISIONS

TOWING

Refer to Figure 7. The nose gear provides

for an applied towing load

not

to exceed 20 000

pounds applied in any direction in a plane paral le l

to the ground. Each main

gear

provides

for

a towing

load

not

to exceed 20 000 pounds. This load may be

applied in any direct ion within 45 degrees

of

a line

parallel to the longitudinal axis

of

the airplane and

wi thin 30 degrees in a plane paral le l to the ground.

The

nose landing gear has lugs for the attachment

of

a tow bar. Lugs are also provided on the main gear

for the attachment of ropes or cables.

The

tow bar attachment lugs

on

the nose landing

gear have been intentionally equipped with bushings

of

a different size from those installed on Constella

tion models in order to require the use

of

a new tow

bar.

Tow

bar shear pin breakage would be a problem

if a Constel la tion tow bar were used

on

the Starliner.

Conversely if the heavier 1649A tow bar were used

on

Constellation models serious damage to the nose

gear attachment structure

might

result.

J KING

Refer to Figure

There are two jacking

points

on

the wing and two

on

the fuselage.

The

fuselage nose jacking point requires the attachment

of

a jacking

pad

assembly.

At

all other wing and

fuselage jacking points the jack pads are inserted in

th e

recep tacles which are integral parts

of th e

structure.

Jack points on the main and nose landing gear are

integral parts

of

the strut assemblies and permit jack

ing in any loading configuration including maximum

gross takeoff weight of 156 000 pounds.

Means are provided for the attachment

of

weights

to the airplane nose to balance the airplane during

overhaul operations such as removal

of

engines.

MOORING

A total

of

six tie-down points are pro

vided. These are shown in Figure

LEVELING

Fifteen leveling alignment

and

symmetry

points are provided see Figure 7 All points are


11/32

• HOIST FIITINGS COMPONENT PARTS

X

MOORIN

POINTS

• TOWIN POINTS

JACK FITIINGS

• LEVElING POINTS

•

\

•

•

•

\

•

4 3

65 2

57 2

imensions

shown are minimum

turning radii

igure

7

round Handling Provisions

•


12/32

ILLUSTRATION

CODE

ENG

OIL

A

ENG OIL B

FUEL

A

FUEL

B

FUEL C

FUEL 0

HYD

OIL A

HYD OIL

B

HYD

OIL C

HYD

OIL 0

HYD

OIL

E

ALCOHOL

WATER

lAC OIL A

lAC

OIL

B

lAC

OIL

C

REPLENISHING

T

CHART

CAPACITY

OF EACH UNIT

SPECIFICATION

UNIT

LOCATION

OF FILLER U.S.

IMP. METRIC

GALS. GALS.

LITERS

AND GRADE

ENG OIL TANKS

TOP

OF

ENGINE

NACELLES

47 50 39 50

179 80 MIL 0 6082A

GRADE

l l20

WAD

5815,

GRADE

120

RESERVE

ENG OIL

TANK

LEFT WING/FUS FILLET

64 00 53 30 242 20

SAME

AS

ABOVE

FUEL TANKS

NO 1

AND

4 NEAR

WING TIP, TOP WING

1350 00

ll24 10

5109 70 MIL F·5572

GRADE

l l5 145

FUEL

TANKS

NO 2

AND

3

BETWEEN NACELLES,

TOP OF 1350 00

ll24 10 5109 70 MIL·F·5572

WING

GRADE

l l5 145

FUEL TANKS

NO 5

AND

6

INNERMOST

TANK

ACCESS

1350 00

ll24 10 5109 70

MIL F 5572

DOOR, TOP

OF

WING

GRADE

l l5 145

FUEL

TANK

NO 7

ADJACENT TO RIGHT

WING/

1500 00 1249 00 5677 50

MIL·F·5572

FUS FILLET, TOP

OF

WING

GRADE

l l5 145

HYD RESERVOIR

NO 1 HYD

SERVICE

AREA

3 60 3.00

13 60 MIL·0 5606

HYD

RESERVOIR NO 2

HYD SERVICE AREA

2 10 1 70

7.90

MIL·0 5606

HYD AUX

RESERVOIR

HYD SERVICE AREA

3 60

3.00 13 60

MIL 0 5606

BRAKE

MASTER

CYL

THROUGH FUS NOSE ACCESS

0 10

0.08

0 38 MIL·0·5606

RESERVOIR

DOOR

HYD

EMERGENCY

TANK

THROUGH FUS NOSE ACCESS

4 50

3.70 17 00

MIL 0 5606

DOOR

ANTI·ICER

TANKS

TOP AFT END NACELLES

20 00

16 60 65 00

MIL F·5566

ALCOHOL

NO 1

AND

4

WATER TANK

FILLERS AND

RIGHT

LOWER

SIDE OF

FUS

60 00

50.00

227 10

PORTABLE WATER

FILLER CONTROLS

CABIN

SUPERCHARGER

NACELLES

NO 1

AND

4 2.30 1 60

7.50

AEROSHELL lAC

HEAT

EXCHANGER

HYD

MOTORS AND FANS,

AEROSHELL lAC

RIMARY

UNIT

AND

LOWER AFT NACELLE

NO 1

0.12

0.10

0 45

SECONDARY UNIT

PRIMARY UNIT AND

LOWER

AFT

NAC

ELLE

NO.

4 0 12 0 10

0 45

AEROSHELL

lAC

SECONDARY

UN

IT

COOLING

TURBINES

LOWER

AFT

NACELLES

NO 1 0 13

O ll

0 49

AEROSH

ELL

lAC

AND

4

Figure 8 Replenishing ata hart

external with the exception

o

an engraved plate

located beneath the cabin floor same as 1049G .

This plate indicates the level

o

the airplane laterally

and longitudinally when used in conjunction with a

plumb bob suspended from the airplane structure

above the plate.

HOISTING

Provisions are incorporated for hoisting

individual major assemblies

o

the airplane

as

shown

in Figure

o

provisions are made for hoisting the

fuselage/wing assembly or the entire airplane.

REPLENISHING

Replenishing data is given in Figure

8.


13/32

ortion o

Nose Gear

ownlock Mechanism

Showing Spring Tube ssemblies

VIEW

ORRE T

IlST LL nON

VIEW

Il ORRE T IlST LL TlON

ONSTELL TIONS

Most cases

of

failure of the

nose gear to lock in the down position have been

caused by

too much dirt

and

not enough lu ric nt

in the downlock mechanism. Thorough periodic

cleaning and lubricating

of

downlock components

s

specified in the maintenance manuals are two of the

best guarantees against malfunction.

Several reports on downlock malfunctions also

verify the need for close inspection and above all a

conscientious

ground oper tion l check

after any

maintenance has been performed on landing gear

components. This will go a long way towards ensur-

ing that all parts have been correctly assembled

installed adjusted and are in good working order.

he

photographs illustrate a case in point where

this was not done. he inner spring tubes of the nose

gear downlock mechanism were incorrectly installed

causing them to bind. hen the downlock failed to

lock. View A of the illustration shows the inner

spring tubes correctly installed.

ote

that the tubes

are not symmetrical and that the longer shoulders on

each tube face towards each other at the point where

the pivot bolt joins the tubes and the link arm. These

shoulders position the inner spring tubes in relation

to the outer tubes so that a smooth telescoping action

is possible. This telescoping action occurs s the

downlock shaft pushes past the downlock and into

the locked position.

ote

also

that

when installed

correctly the two spring tube assemblies appear to

be parallel.

View B shows the inner spring tubes installed

incorrectly. ote that the longer shoulders on the

inner tubes face away from each other and that the

sides of the inner tubes are now forced against the

outer tubes s the whole assembly is out of alignment

and the two tube assemblies are not parallel.

he inset macrograph

of

a tube which was installed

incorrectly reveals the chafing which occurred until

finally the assembly seized and the gear could not

lock down.

o prevent nose gear downlock problems:

Clean and lubricate in accordance with the

instructions in the maintenance manuals. Main-

tain and assemble gear components with extreme

care do not overtighten nuts on any pivot

bolts in landing gear moving linkage then

always inspect and ground check the affected

gear for correct operation. A A


14/32

We

selected the Scovill adjustable fastener to cor

rect this condition because it incorporates a feature

which compensates for wear

by

providing a sheet

take

up of

approximately .100 in. between the access

door and the door jamb.

When

this fastener is used

properly, it provides a firm door-to-structure attach

ment and reduces motion and wear.

CONSTRUCTION

As shown in Figure

the Scovill

fastener consists

of

a floating-type receptacle assembly

and a stud

and

grommet assembly.

The receptacle assembly consists

of

an inner

member which screws into an outer-member. The

outer-member floats in a housing which is attached

to the door jamb structure.

The

inner-member consists

of

an externally threaded sleeve which incorporates a

spring and a special de tent washer (see Figure

2 .

OPERATION

To

engage and adjust the fastener, a

hand-driven

No.

2 Phillips screw driver is used to

press the stud assembly into the receptacle inner

member ( slot ted to admi t the stud and to turn the

stud clockwise. As the stud turns, the cam lugs on

the stud engage shoulders in the counterbore

of

the

inner-member

and

screw the inner-member into the

receptacle outer-member.

The

cam lugs also engage

covill diustable s

Figure 1 Photo of

Scovill

Receptacle

Retaining

Ring

and

Stud and Grommet ssembly

1 49A G H

The

use

of

Airloc fasteners on the

left

and right inboard oil tank access doors (lower surface

of

wing leading edge has been discontinued. Scovill

fasteners are being used instead, effective as follows:

1049A, Serials 4407 through 4412 and 4433 and up;

1049G, 4642 and 4645 and up; 1049H, 4801 and up.

The part numbers for the doors with Airloc fasteners

are LAC

PIN

312629-3

LIR or

PIN

312629 500

L/R

The modified door assembly with Scovill fasten

ers has been redesignated

as

LAC

PIN 312629 5 1

for the lef t wing and -502 for the right wing.

In addition to the change in fasteners, the door

cutout in the wing has been reinforced and a more

durable nylon chafing strip was installed on the door

jamb.

These changes were made because some operators

have reported difficulty in keeping the door assembly

securely attached and in a few instances, a door has

been lost in flight. Operators attributed the problem

to excessive wear

of

the mating surfaces around the

door s perimeter. In turn, the wear was attributed to

the chafing which resulted because it was not always

possible to maintain a firm attachment between the

door and the door jamb.

12


15/32

notches on the mating face

of

the special

spnng-

loaded detent washer. The action

of

the detent

washer, plus the tension on the stud afte r it has been

tightened, prevents the stud from loosening.

The

detent washer has two tongues

on

the outer perimeter

which engage grooves in the inner-member and so

prevent the detent washer from turning.

As a further precaution against the fastener loosen

ing in flight, the threads on the inner- and outer

members are lubricatedwith dry-film lubricant during

manufacture so that even though the fasteners become

saturated with oil vapor or l iquid during service, the

torque required to turn the inner-member should

remain near the design optimum.

When

installing t he acces s

door, if

turning the stud

clockwise seems to force the stud (and the access

door) away from the receptacle, it means that the

receptacle is out of adjustment;

that

is, the inner

member is screwed into the outer-member and away

from the stud too far

for

the cam lugs

on

the stud to

engage the shoulders of the inner-member.

To

correct

this, press the stud firmly into the receptacle and

unscrew the stud

and

inner-member one or two ful l

turns

counterclockwise.

This will screw the inner

member closer to the cam

lug

end

of

the stud. Then

press the stud in and turn it clockwise (approximately

one-quarter turn) until you feel the stud engage the

shoulders in the inner-member. Continue to turn the

stud (and inner-member) until the fastener

is

tight.

In addit ion to sheet take up

in.)

provided

by

the inner-member screwing into the outer-member,

the vendor is studying ways to improve the stud and

receptacle to provide an addit ional 6-in.

of

sheet

take up. These changes will also provide the mechanic

a more positive feel of stud engagement and will

practically eliminate the need for backing off the

inner-member to engage the stud (as described in the

preceding paragraph). The vendor has indicated that

fasteners incorporating the additional improvements

will be available in the near future.

Tighten the studs approximately the same

amount

as a

standard /4

-in.

screw.

Use a

No

Phil/ips

hand-

driven screw

driver only

To

avoid damaging the

recess

in the

head

of the

stud,

do

no t use

a Reed

and

Prince

screw driver

or

a power screw driver of an y

kind

To

loosen the

stud

and

open the

access

door,

approximately one-quarter

turn

in a counterclockwise

direction is required. However, to faci litate the next

fastening,

turn

the stud

(and

inner-member) approxi

mately one more turn counterclockwise. This wil l

unscrew the inner-member sufficiently to make it

easier to feel that there is positive engagement

RECEPTACLE

INNER-MEMBER

RECEPTACLE

OUTER-MEMBER

(Floots

in

housing)

STUD

SPRING-LOADED

DETENT

WASHER

(Tongue

groove

prevents rotation

DOOR

JAMB

STRUCTURE

NYLON

CHAFING

STRIP

In

sections A-A ond

B-B

stud an d

r ivet are shown

solid; not cut

by

section

plones.

Stud

is

shown

in

locked

position

in

both views.

/1

- -

I I

rG:ll

rl-\:7ll

B

I I

B

t

I

/

-L./

Figure 2

Two Cutaway

Sections Show

How

Cam

lugs

on Stud Engage nner· emberand Thread tinto Outer Member to Provide

Sheet

Take

Up

Feature

13


16/32

between the cam lugs on the stud and the shoulders

on the inner-member when the access door is

reinstalled.

INSTRUCTION PL C RD

The

modified access door

LAC P 312629-501 and -502 included a placard

which provided instructions for opera tion

of

the

Scovill fastener. However, the placard has been

moved to a more conspicuous location

and

the

instructions

on

the placard have been revised to

include greater detail see illustrat ion .

The new placard will be installed on production

~ i r p l n s as soon s possible. Commercial operators

whose aircraft were delivered with the Scovill fasten

ers in the modified door assemblies may procure the

new placard LAC

P

497750-1 through our Com

mercial Spares Department.

SERVICE ULLETIN Service Bulletin 1049 SB 2851

is

now being prepared for commercial operators

who

wish to incorporate the production modification and

will include information regarding the installation

of

the Scovill fasteners, reinforcement of the door jamb

structure, and installation of the more durable chafing

material. This Service Bulletin

1049 SB 2851

is

scheduled to be transmitted approximately June 22,

1957 and the letter

of

transmittal will include infor

mation regarding procurement of parts.

The fasteners are manufactured y the Scovill Mfg.

Co. of Waterbury 20, Connecticut. A A

D•C

enerator

Overvolt

J •

49

S RI S Occasionally the overvoltage relay panel

LAC

P

617167-3, General Electric

P

CR 2781

M146D

is

reported to trip for no apparent reason.

The

trouble may be that the internal shipping pads

which support the shock mounted relay during trans

portation, have been lef t inside the panel assembly

see illustration . the shipping pads usually cor

rugated cardboard were not removed before the

panel was installed, nuisance tripping of the relay

may result because the pads defeat the shockmount

ing protection.

The illustration shows the notice which is fastened

to each panel calling attention to the shipping pads

inside. t is necessary to remove one screw shown in

the photograph slide the cover off the panel assem

bly, and then very carefully remove the four shipping

pads wedged under the relay.

The

cover can then be

replaced and the unit installed.

heck

fo r

shipping

pads

in

overvoltage

relay panels which trip for no apparent reason.

Inspect all panels for shipping pads at any major

overhaul period. Be sure to remove pads when

installing a new unit. A A


17/32

1049C, E G, H The CAA-approved airplane

flight m anua ls for the subject Super Constellation

models restrict the maximum allowable fuel in each

tip tank Tank Nos. 1A and 4A to 1200 pounds

during landing. Lockheed has imposed this structural

limitation because of the high loads placed on the

wing during a landing in which each tip tank contains

more than the 1200-pound maximum fuel load. This

restriction presents no hardship under normal oper

ating conditions because fuel management procedures

recommend t ha t tip tank fuel be consumed during

initial cruise flight. In addition tip tank fuel can e

dumped in flight should an emergency arise.

The

CAA has requested t he inst all at ion

of

ade

quate placards to advise flight crews

of

this.landing

Landing

Limitations

With

p

Tanks

limitation.

To

comply with this request we are now

placing a placard containing the pertinent informa

tion on the flight engineer s middle instrument panel.

Production installation of the placard will be effe -

tive on LAC Models and Serials: 1049G 4672 an d

up; 1049H 4813 and up. Lockheed will install the

placard on commercial aircraft which are delivered

with tip tanks installed. tip tanks are to be installed

after delivery of t he aircraft th e pla card will be

included in the tip tank kits.

Operat ors w ho have purchased Service Bulle tin

10491SB 2552

and have installed tip tank provisions

on their in-service aircraft may procure the new

placard LAC

P N

496910-1 through our Commercial

Spares Department. A A

Installation of Leather ack UpRing

n rakeSelector Valve

10498ASIC TH ROUGH H During overhaul of brake

selector valves LAC

P N

548350 or P N 548350 0 1

some mechanics have found it necessary to stretch the

l ea ther back-up ring

P N

AN6246-5 in order to

i nsta ll the ring on the actuat ing piston

P N

Bendix

1006199 or P N 1006199 2.

Stretching the back-up ring increases its diameter

and i t tends to lap over and wedge between th e edge

of the piston groove and the valve body after the

piston and ring are installed in the selector valve

body. This causes the piston to tilt slightly and allows

hydraulic fluid to leak out of the valve past the cam

shaft opening.

The Pacific Division of Bendix Aviation Corp.

vendor of

the brake selector valve recommends the

following two methods which may be used to install

the leather back-up ring so that it will not be stretched

permanently out

of

shape.

• The first method is to soak the back-up ring in

w arm wat er and install it i n t he piston groove. o

not install the O ring at this time.

Bake the piston

with back-up ring installed for 5 to 7 minutes at

300

0

P 149°C . The back-up ring should return to

PISTON

SSEMBLY

Typical

BR KE SELECTOR V LVE

For clarity, some parts are

omitted from

this

exploded

view.

its original shape. After the piston cools install a

new a-ring of the proper size.

• Another method

is

to coat back-up ring with hydrau-

li oil to aid in sliding it over the piston. A A

15


18/32

LL

ONSTELL TIONS As the onstellation

and

Super Co nstellatio n h ave g ro wn in size and weigh t

through the years more hydraulically operated equip

men t h as been added. Thi s increased demand on t he

hydraulic system may combine with other conditions

to cause variations in nose wheel steering action.

Occasionally a pi lo t may feel t ha t a pa rticular air

p la ne steers h ar de r o r more slowly tha n others.

He

may even feel that the steering action in the same

a irp la ne varies f rom time to time and may ask the

mechanic to check over the system to get a better

steering response.

There

are cer tain adjustmen ts and changes to t he

nose ste ering system which will aid in eliminating

these problems

if

they appear on a particular air

plane. First we will explain the variations in operat

ing conditions affecting nose gear steering. Then we

will describe the adjustments which will ensure that

the nose steering system operates at its best efficiency.

NORMAL

OPER TING H R TERISTI S

uring

taxiing or landing roll-out the output of the pumps

in t he secondary hyd raulic system s relatively low

because of the low rpm of the engines driving them.

Yet the wing flaps brakes heat exchanger fan and

th e nose steer in g may be i n simultaneous operatio n

and the demand on the hydraulic system s quite high

at this time. These two o pp osin g factors o ft en cause

t he variations in nose wheel steering action which

are noticeable to the pilot. For example under these

cond itions t here may be times when the p il ot s nose

st eerin g contr ol wheel wil l app ear to be h ar d to t ur n

simply because he s att empt in g to t ur n i t faster t han

available system pr essur e can tu rn the nose wheel.

Or

t he tr ou bl e may be an actual malf un cti on which

must be f ou nd y systematically trouble shooting the

nose steering system.

TROU LE

SHOOTING

The first requirement in normal trouble shooting

s to check all functional parts in the steering system

for adjustment lubrication and operation in accord

ance with the applicable maintenance manual.

In this regard it s i mp or tant f or th e p il ot and the

maintenance man to agree on the difference between

hard steering and slow steering. Hard steering may

be defined as t he condition when th e p il ot s st eer ing

control wheel s difficult to rotate

y

h an d a t a norm al

rate of turn. Slow steering may be defined as a slower

than-normal response of the nose wheel to the normal

movement

of

the steering control wheel. Each of

these two types of steering action can affect the other

so tha t one

s

o ft en conf used wit h t he oth er . But we

can resolve most trouble shooting problems if we

make use of the following general statement which

s illustrated y Figure

Most

of the

mal functions which cause h rd

steering will occur the region shown s Area A

in Figure

7.

This region pertains to

the nose

steer-

ing control valve nd its control system Con

versely most of the malfunctions which cause slow

steering will occur the region shown s Area

Figure

re

pertains

to the

hydraulic nd

mechanical portions of the steering mechanism

between

the

control valve

nd

the ground

Of

course t here are exceptions to this b ro ad rule

b ut i t sho ul d h el p th e maintenance man to n ar ro w t he

search f or a mal fu ncti on aft er t he p il ot has carefu lly

described the undesirable symptoms.

HARD

STEERING This difficulty could result from

sticking pulleys o r im prope r ri ggi ng o n the cables

connecting the wheel to the nose steering control

valve on t he nose strut. Stiff ope ra tion of th e nose

steering c ontrol valve due to inte rna l binding may


19/32

igure 2 Removal of

Nose

Steering Restrictor

Valve

I

I

-

I

VIEW LOOKING FORWARD IN NOSE WHEEl WElL

AT

FUSElAGE STATION 195.

Removed ports

are

s ho wn i n dotted

outline.

Replacement

part

is shown

by

white area.

Existing

parts are outl ined by

solid line_

PIN

329612_

46

J piN 667070

T U ~ E

ASSEMBLY

NOSE

STEERIN

4

REMOVED) RESTRICTOR

VALVE

NEW TUBE

ASSEMBLY

REMOVED)

PIN 474925-121 _ L _ n \ ~ ~ = ~

I r ~ J I I

NOSE

------. L


20/32

NOSE L NDING GE R

P RT ll

Y

RETR TED

Nose steering shut off valve

must

close when nose

londing

geor

is within the

tolerance shown

Measure

in straight l ine between

upl ock lug on no se st rut

ond l ow er f oc e of

up lock lotch g uid e.

LINK

SSEMBLY

p

8754

DIT Il

A

igure djustment

of

Nose

Steering Shut·off Valve

New alternate method uses

linear

measurement

POSITION

OF

NOSE

STEERING SHUT·OFJ

V lVE

UNK GE WITH

NOSE

GE R FUUY EXTENDID

W RNIN

While

the nose gear

is

held in the partially

retracted position by hydraulic pressure,

place proper supports beneath the lower

end

of

the strut to prevent downward move

ment of the gear.

ke t s measurement first

If no t c or re ct , e xt e nd n os e geor fully

o nd o dj us t l en gt h o f

link

ossembly piN

28754

see Detoil

A .

retract t he nose gea r slowly until th e tires have just

cleared t he jack and the g ro und a t t he b ottom

of

the

swing and the gear is on the way u

4.

When

the nose gear

is

moving u and aft,

check to see that the steering collar, torque links,

and tires will be free to move throu gh t he full arc of

steering travel. Th en while the gear

is

retracting,

actuate the steering mechanism.

Do

this

by

turning

the pilot s steering wheel back and forth continuously

through an arc which includes at least 45 degrees each

side of the centered position.

5. Stop retracting the gear when the steering cylin

ders no l onge r respond to movements

of

the pilot s

steering wheel. Then release the steering wheel. The

nose steering mechanism should slowly center itself

and return the nose wheels to neutral. Hold the nose

strut in position at this p oi nt by relieving hydraulic

pressure or bypassing flow at the gig so that the gear

does not move either

up

or down.

Check the adjustment of the nose steering shut-off

valve P 466345 either in accordance wi th the

applicable maintenance manual or

as

described below.

Proper adjustment

of

the point at which the valve

shuts off during retraction

is

essential to ensure that

wh en th e nose lan din g g ear

is

completely extended,

the steering shut-off valve

is

completely open and

does not cause additional restriction to the fluid

flow.

ADJUSTING SHUT OFF VALVE A simple alternate

method

of

adjustment to that presented in the main

tenance manual

is

shown in Figure 3. Check to see

that

all l an di ng g ear down lock pins are installed.

Close the m in landing gear manual shut-off valves

in order to isolate the main gear retraction mechanism

from hydraulic system pressure. To adjust the shut-off

valve, proceed with the following steps in sequence:

OT

NOSE STEERING SHUT OFF VALVE

It is

unnecessary to relieve air pressure in

the nose gea r shock strut to perform any of

the steps in this check.

desired, the replacement tube assembly may be

fabricated

by

t he ope ra tor from 308

lis

H corrosion

resistant steel tube of Y2-in. outside diameter and .028

in. wall thickness.

Super Constellation aircraft, LAC Models and

Serials 1049G 4664 and subsequent, and 1049H 4803

and subsequent will be delivered with this change

incorporated.

Before perform ing a fleetwide modification by

removing this valve on Model 749A Constellations

or earlier aircraft, it is advisable to test one or more

aircraft modified in this manner to determine the

reaction of the pilots.

1. Jack t he compl ete a ircraft i n accordance wi th

instructions in the maintenance manual. Remember

that the nose strut swings

ownw r as

well

as

back

w ard during t he first

part

of the retraction

cycle so

jack to ample height for the nose tires to clear the

ground a t t he bot tom of t he swing.

2. Remove the safety pin from the nose gear

downlock. Connect the nose strut torque links. Station

a man in th e cockpit to actuate the steering control

wheel and the landing gear selector handle and to

act as safety man.

3.

Apply hydraulic pressure to the system

by

exter

nal means such as a porta bl e test gig. Place t he land

ing gear selector handle in the UP position and

18


21/32

Do not place the landing gear selector

handle in NEUTRAL. Doing this will

merely connect both sides of the nose gear

actuating cylinder to the return side of the

hydraulic system, and the gear will free fall

if the.supports are removed.

6 See

Figure 3 Measure between the lower face

of the uplock latch guide P N 272558 and the upper

most surface

of

the uplock lug

P

282015

on

the

nose strut. As shown in Figure

3

the distance must

be 41.75

+5.0, 0 .0

in. in a straight line between

these two points.

The

partially folded drag links will

interfere with ordinary m ~ u r i n g methods, however.

To avoid this interference, use a long piece of stiff

wire to form an offset measuring rod. Bend each end

of the wire so that it touches one side of the two

points described above, and bend the middle portion

of the wire to clear the right or left sides of the inter

fering drag links. The wire can then be placed flat on

the floor and the measurement made in a straight line

between its ends, using a steel tape.

An alternate measuring device can be made from

a pair of long wooden laths bolted together like

scissors. This device will serve as a type

of

inside

caliper to make this measurement.

7 the shut-off valve closes outside of the speci

fied range of strut movement, adjust the link assembly

P

287540.

OT

Shorten the link assembly to get an earlier

shut-off point during retraction. Lengthen

the link assembly to make the shut-off occur

later. One-half turn of the clevis on the link

assembly

is

equivalent to approximately one

inch

of

nose landing gear movement meas

ured between the points described above.

Detail A in Figure 3 shows the minimum

allowable length of the link assembly,

which shall be 5.06 in. after adjustment.

The inset shows the position of the shut-off

linkage when the nose landing gear

is

fully

extended.

It should be noted that this method

of

checking

the shut-off valve adjustment does not change the

setting of the valve given in the maintenance manual;

it

is

merely another way to arrive

at

the same setting.

Following the instructions in this article should

ensure that the fluid flow

is

not restricted and that the

best steering rate

is

achieved. A A

Integral

Fuel

Tanks From

orrosion

y

Vern

Dress Lockheed materials

and

pwcesses

development

engineer

49

S RI S Two years ago the Field Service Digest

Vol. 1 No.6 carried a comprehensive report on

the control of corrosion. This report pointed out

areas most susceptible to corrosive attack and recom

mended suitable maintenance procedures. Service

Bulletin 1049/SB 2630 dated March 20, 1956

describes in greater detail the action necessary to

remove existing corrosion in integral fuel tanks for

commercial models. For military models, similar

information is included in the applicable Structural

Repair Manuals. As additional protection, these docu

ments also include irrstructions for adding more drain

holes in the stiffeners and for extending the area

protected with sealant.

Because the inroads

of

corrosion in the wing fuel

tanks can cause serious structural damage, we should

like to focus attention again on the preventive meas-

ures necessary for the treatment

of

milled skins in

this particular area. the inspection and preventive

maintenance outlined herein are given timely and

thorough attention, corrosive attack may be prevented.

It

is

possible that through neglect, structural damage

requiring costly, major repair might eventually occur

with no external evidence

of

the corrosion.

INTEGR LLY STIFFENED STRU TURE

By using one integrally stiffened panel instead of

an assembly of a great number of smaller parts,

designers have eliminated many

of

the complications

and weaknesses inherent in bits-and-pieces construc

tion, thus greatly increasing strength and yet decreas

ing weight and fuel leak problems. Integral structure

has proved so efficient

that

it has been widely adopted

in the airframe industry.

9


22/32

When tanks we re constructed mostly from clad

alloys, no special precautions against corrosion were

necessary. The concept of integral construction, where

cladding can not be used, has changed the picture.

The best information available when the first 1049

airpl anes we re be ing bui lt showed t ha t t he Buna -N

fill-and-drain protective coating LAC 1-781 Type

II

woul d be a de qua te to prot ect t he newly developed

i nt egral s truct ure i n t he fuel tanks from corrosion.

Indeed, many 1049 s have been in service for over

four years without evidence of corrosion in the fuel

tanks. In general, these airplanes use only fuel which

has been produced in accordance with strict U.S. MIL

Specifications. For this reason, we suspect that the

corrosion which has been encountered

so

fa r may be

caused to a great extent by acids and corrosive sub

stances which t he wa te r collects from ot he r fuels.

These contaminants apparently penetrate the Buna-N

film and attack the milled skin.

We

do not know of

any instances wh er e corrosion has occurred under

sealant which was properly applied and properly

coated with Buna-N.

INSP TION

the integral fuel tanks have not received the

additional protective coat of sealant in accordance

w ith the inten t of Service Bulletin 2630

or

the appli

cable military Structural Repair Manual, or if the fu el

tanks have not been inspected recently for corrosion,

the condition of the milled skin panels in the wi ng

fuel tanks s ho uld be as certained

by

careful inspec

tion. As shown

by

the crosshatched zones in Figure 1

the areas which s ho uld be examined w ith p ar ticu lar

care are as follows: lower surfaces of the surge boxes;

an area for a distance of about

2

inches outboard

from the bulkheads at WS 87,

WS

215, and

WS

480,

in particular; each access door and surrounding area.

The Buna-N coating must be removed for this pur

pose, since it may become opaque after long exposure

to moisture and hide corros ion that may be present.

Methyl Ethyl K eton e MEK should be used to

soften the Buna-N protective coating. Then remove

the Bun a- N w ith scrapers made f ro m micarta, hard

wood, or red fiber.

To

avoid the necessity of a tank

WS

191

o

WS

458

t

OUTB O

NACELLE

i

Stringer No. 24

Stringer No.

28

)

S tr in ge r N o. 1

Stringer

No.7

S tringer No. 13

S tringer No. 18

WS WS

215

239

INB O

NACElLE

WS WS

87 105

1

;

L ower s urface of right inner wing s hown. Left wing

is

opposite.

W

UTl t

WS

463

Stringer No.7

--Stringer No. 10

--Stringer N o. 12

WS

668

Lower s urface of right outer wing s hown. Left wing

is

opposite.

Figure 1

Crosshatching

Indicates Areas in

Fuel

Tanks

Where Additional

Protection Is

Required


23/32

soak check, be careful not to remove more sealing

compound than is necessary, particularly from areas

close to joints.

X MPL S

In

looking for corrosion, itmust be remem

bered that corrosion may be present, but not recog

nized as such. Therefore, it

is

advisable to check

carefully any areas which have an abnormal appear

ance which could conceivably be attributed to

corrOSIOn

corrosion

is

present, it may have a variety

of

appearances as shown in Figures 2 through 4 The

figure titles describe the corrosion and where it was

found in the fuel tank. In some cases it may appear

as isolated spots of discoloration which can easily be

removed by light sanding. Where the ends

of

metal

grains are exposed, such as at cutter run-outs at pads

or in a-ring grooves in fuel tank access doors ,

corrosion may appear

as

dark lines separating layers

of metal as shown in Figure 2 In other cases con

siderable areas may be discolored by dark patches of

powdery corrosion having a salt and pepper appear

ance

as

shown in Figure

3

Occasionally the attack concentrates in very limited

areas forming deep pits as shown in the micrograph

in Figure 5 Sometimes blisters are formed,

as

shown

in Figure 6

RE OMMEND TIONS

To

combat this corrosion problem, on future pro

duction airplanes, we have extended the application

of

sealant compound to include low areas in the fuel

tank where water can collect see Figure

1 .

This

was accomplished on the following LAC Models and

Serials prior to delivery:

1049A Serial 4634 and subsequent

1049B Serial 4171 and subsequent

1049G Serial 4620 and subsequent

1049H Serial 4801 and subsequent

Service Bulletin 1049/SB-2630 or the applicable

military Structural Repair Manual describes in detail

the anti-corrosion precautions which should be taken

on in-service aircraft. In these documents we recom

mended

that

quick repair brush sealant LAC 1-766

Type III Products Research PR 5401K , or PR

1221BT-FAST, be used as a cover coat to extend the

protective barrier against corrosion. This sealant is

impervious to water and the corrosive contaminants

it may contain. A new and better sealant, LAC 1-778

Type II Products Research PR 5701K or PR 1422

BT-FAST which have improved resistance to fuel and

Figure 2 Enlarged

View

of Fuel Tank Access Door Illustrating

Edge·Grain

Attack in

D·ring

Groove

AI and Blisters Bl

contain a chromate corrosion inhibitor are now avail

able. neither of these is available, standard brush

sealant LAC 1-775 Type II

PR 7101K ,

or

PR

1221BT, may be used.

Service Bulletin 1049/SB-2630 is now being revised

to specify the improved sealant protective coating and

LAC 1-775 Type I fillet sealant to fill up the low

areas in the fuel tank sufficiently so that all water will

drain to the sump drain valve. Future production

aircraft are also scheduled to incorporate these

improvements. The

revised Bulletin

is

scheduled to

be transmitted by the latter par t of June, 1957.

Continued on next page

Figure

3

Corroded

Area

Near

Bulkhead in No 1 Tank

at

215

Note plugged drain holes


24/32

figure 4

Corrosion

Along Stiffener at Access Door

Cutout

No

Tank

Attack was 3B in

deep in fillet

figure

5

Micrograph Section

Through Typical

Pit

x2

GENER L

To protect the fuel tank structure i t

is

most important that the procedures recommended

below be followed carefully and in proper sequence.

Essentially the treatment consists

of

the following

steps:

• Inspect fuel tank interiors.

• Remove corrosion

i f

any

is

found .

• Apply a chemical film.

• Coat corroded areas and low areas with sealant.

• Coat the sea lant with Buna-N.

any step

is

slighted

or

done

out of

turn the

treatment is useless since each material must adhere

to and protect the coat beneath it. Thus

if

the sealing

compound

is

not coated with Buna-N the fuel may

eventually leach

out

the rubber content

of

the sealing

compound so that

water

can penetrate to the structure

and attack the chemical film. Since the chemical film

must be slightly soluble to provide the chemical action

necessary to neutralize the corrosive effect

of

water

long contact with water may eventually dissolve the

film and the metal panel wil l begin to corrode. The

successive barriers against fuel and water must there

fore be carefully constructed

and

maintained.

R MOVIN

ORROSION

The

first step

is

to use

No. 280 abrasive cloth

or paper

to sand off all affected

areas and remove

as

much corrosion as possible see

Figure

7 .

deep pits or lines of corrosion are left

an air motor driving a rubber wheel impregnated with

aluminum oxide abrasive may be used to remove

enough metal to get to the far thes t extent

of

attack.

We

have previously recommended that power tools

not be used since they may remove too much metal

and care should be taken in this respect. Also to avoid

int roducing dissimilar metal into the area do

not

use steel wool or wire brushes.

•

. :

;

X2

figure

6 Micrograph Section Through Blister Showing

Exfoliation

frequently Encountered x2

22


25/32

Where

blisters have formed or corrosion occurs at

the ends of metal grains, the corrosion attack fre

quently shows a preference for traveling alon g the

grain or direction of rolling or extruding for con

siderable distances. This results in a delaminating

or

exfoliating effect

as

shown in Figure 5 and Figure

8.

In order to ensure that the end of the attack or bottom

of the pit has been reached, the abraded area should

be etched

by

swabbing with a 5 to

10

per cent solution

of phosphoric acid and allowing it to stand for

11;1

to

2

hours.

corrosion is still present, it will show

as

a dark line or spot and additional metal must be

removed.

After the etch test has shown t ha t all corrosion

has been removed, the depth of metal removal should

be measured and structural disposition made. The

applicable Structural Repair Manual gives the damage

limits which must be observed in removing metal

from integrally stiffened panels.

Should a customer encounter corrosion damage not

subject to the typical repairs shown in the Manual,

he is invited to write to our Field Service and Training

Division in the usual manner.

R INISHING

After al l the c or ro si on p ro du ct s h av e b ee n

removed, blend and smooth the metal surface with

No.

400 abrasive cloth. phosphori c aci d etch has

b ee n us ed, b e s ur e that the last trace of it has been

removed by rinsing area repeatedly with water and

drying with clean wipi ng cloths. Then clean the

affected area with MEK or ethyl acetate. Check that

all drain holes in stiffeners are open.

2 Brush on Spec MIL-C-5541 c hem ic al film

solution Iridite 14-2

or

Alodine 600 or 1200, or

e luivalent . After 5 m in ut es , b lo t t he s ur fa ce dry

WIth clean, lint-free cloths or tissues. Do not rinse

with water, as water will dissolve the fresh chemical

film. further cleaning is necessary use MEK

or

ethyl acetate.

3

Apply t wo b ru sh co ats of LAC 1-778 Type

II

sealing compound over the reworked area and

for

approximately another 6 inches around it. The

total thickness

of the

t wo coats should be approx

imately

1/32

in. Allow 2 hours drying time a t

77°F

25°C

with relative humidity of 50 per cent

bet ween coats. Be car ef ul

not

to plug dr ai n hol es

in

the stiffeners.

Instructions

for

mixing and applying sealant can

be found in

Sealing

Constellation Integral uel

Tanks Vol.

1,

No 4

of the Field Service Digest

dated January-February 1955. These instructions

should

be

followed to

the Ie

er to ensure a

good

job of sealing.

4.

When the

se alin g c om po un d is

no

longer

tacky this t akes approxi mately 4 hours at 77°F

25°C with

relative humidity of 50

per cen t ,

Figure

7

Appearance

After

light Sanding

of aCorroded

Area

Similar

to Figure 3. Pits .010 to .020

in. deep

may be found

and

in cutter run-out

may

follow

grain for

.25 in . or more.

Figure

8

Macrograph

Section

Through Access Door Shown in Figure

2

Shows Attack

Along

Grain

End

AI

and

Blister IBI xlO

brush on t wo coats of LAC 1-781 Type I Buna-N

protective coating.

Air

dry t he f irst coat 5 minutes

before applying the second coat. All ow at le as t 1

hour a ir d ry after

the

second coat before refueling.

5

Make certain that all of the drain holes

the

structure are

open

so

that

moisture will

drain

to

t he s um p d ra in v alve.

6. t he c ro ss ha tc he d a re as s ho wn

in

Figure 1

have not received the addi ti onal protect ive l ayer

of se alant c om po un d, u se MEK to re move the

Buna-N

and

ac hie ve a c om pl et el y c lea n s urfa ce .

Inspect for corrosion,

and

remove any found. Then

apply a protect ive coat of sealant to these areas in

accordance wit h steps 3

through

5 above.


26/32

649

749 749A 1 49BASIC THROUGH H 649A

The AiResearch Manufacturing Company has recently

released Service

Information

Letters No. 100-72

and

100-74. These letters contain information on oil

foaming which will be

of

interest to all operators of

Constellation and Starliner aircraft equipped with

AiResearch cabin superchargers.

In essence AiResearch SIL 100-72 recommends

that no silicone oils or greases be used during the

overhaul and assembly of cabin superchargers which

use Aeroshell

lAC

as

the system fluid.

In one particular instance a few drops

of

DC-200

silicone oil were used to lubricate a rings and some

moving parts during assembly of a supercharger.

When the unit was placed in operation the silicone

oil caused the system fluid Aeroshell lAC to foam

and the fluid was lost through the vent line. Tests

have indicated that the greater the concentration of

silicone oil the quicker the system fluid will foam.

Additional investigation of the foaming problem

led AiResearch to issue SIL 100-74 which states that

foaming can be caused by the presence of residual

petroleum lubricating oil mixed with corrosion pre

ventive compound. This letter recommends

that

Aeroshell lAC be used to flush all cabin super

charger oil coolers and oil temperature regulators

overhauled or manufactured by AiResearch prior to

December 1956 before they are placed in service.

Operators

who perfo rm their own overhaul of

cabin supercharger oil coolers and temperature regu

lators should flush these components with the oil

Using the wrong lubricants

during overhaul

causes

Oil Foaming

•

In

iResearch

abin

uperch rgers

they normally use in their cabin supercharger system

before installing or storing them. In any case the

preservative oil Specification MIL-C-6529A Type 3

can cause foaming and should not be used for flush

ing unless the components are to be stored in a

severe environment where extra protection from cor

ro i n necessary.

An additional recommendation in AiResearch SIL

100-72 declares that no lubricant containing molyb

denum disulphide in any form should be used during

the overhaul and assembly

of

supercharger compo

nents. The undesirable effects of molybdenum disul

phide in the supercharger system were noted during

development of the sprag-type clutch used in some

AiResearch cabin superchargers.

The

addition of

molybdenum disulphide to the lubricant so reduced

friction between the sprags and the races that the

sprags slipped when the clutch was engaged. Under

high loads this condition would eventually result in

failure of the sprag clutch assembly.

t

is believed

also

that

the same effect can occur with the roller

clutch used in some models

of

the subject cabin

superchargers.

Both AiResearch letters serve to emphasize the fact

that the mixing of lubricants or system fluids is not a

good policy and the results may be unpredictable.

During

ny m inten nce

operations on cabin super-

charger or

on

the

components

o a cabin supercharger

system

in

which Aeroshell AC used particular care

must be t ken that the system fluid not diluted or adulter-

ted by ny

other fluid or substance


27/32

Wing

lap

Short ink Rollers

/ 0 SHORT LINK

Under some load conditions, the high drag cre

ated

by

the sleeve rotating in the short link will

cause the seized roller assembly to skid in the flap

track. The skidding action will create flat spots on

the roller and may result in jamming of

the

mechanism.

To

remedy this situation, it

is

recommended that

the AN31O-6 nut be tightened no more than ing r

ti ht

as specified

on

the engineering drawing. This

information will be included in the applicable main

tenance manuals at the earliest possible date. .. . . ..

a;:; =---

USHING

W SHER

ET IL

A

WAS HER: l \ \ l l

I T T lNG

Details shown

are

for Models 1O 49A B and

C

Other

models may vary slightly.

fl P

Reference

ONSTELL TIONS

Several reports have been

received concerning the excessive wear

of

the wing

flap short link rollers see illustration . Investigation

of this problem revealed that excessive wear is caused

by

the following conditions:

When

the AN31O-6 nut on the short link roller

is

overtightened, the bearing may be preloaded to

such a degree as to cause brinelling, which in turn

will probably cause the bearing to seize. Because the

outer race of the bearing is cemented to the sleeve,

seizing

of

the bearing will cause the sleeve to rotate

in the short link during flap operation.

C RRI GE SSEMBLY


28/32

WARNING

LAC

and

WAD do o approve the practice

of insta ll ing oil def lector assemblies

during

engine

overhaul on

any

engines which

will

be operated under

the following conditions:

1. When the using aircraf t employs the

six generator electrical system

2. When the right hand generator drive

of the

engine

might be used to operate any

accessory other than the cabin supercharger

After the LAC oil deflector assembly has been

installed, the accessory drive cover and a new gasket

should be reinstalled and the

nut

properly torqued.

Thi s cover should not be removed unless t he engine

is

used in an out board position on the aircraft, and

the right hand accessory drive is required to operate

the cabin supercharger. such is the case, the acces-

sory drive cover should not be removed until the

cabin supercharger shaft disconnect assembly is to

be installed.

During overhaul of engines not restricted by the

two notations above, the

WAD

P

428123 oil seal

assembly should be removed from the righ t hand

generator drive pad of the engi ne and

if

serviceable,

retained as a spares replacement for future use. Next,

the LAC oil deflector assembly P 491490-1, or

the superseded assembly LAC P 470186-3, should

be i ns tal led in t he same location on the engine. LAC

P 491490-1 deflector assembly is interchange

able wi th the superseded

P

470186-3, but pro

vides improved lubrication because

of

closer design

tolerances.

p i _ l l l

IlIIlKl1lI

ASSElIIlY

p

4914lO-1Dollod,,_iIy

isinl

P


29/32

dded Hole in Engine Upper owl Panel Makes lubrication Easy

49

SERIES

The operating linkage for the carbu

retor ram air and alternate air doors requires periodic

lubrication. The forward lube fitting on the ram air

door link is not easily accessible on some early ver

sions

of

the upper cowl panel equipped with ram air

doors because a screw-fastened access panel must

first be removed. Consequently this fitting may some

times be neglected during lubrication.

A I-inch diameter hole dri lled through the

access

panel in the location shown in the illustration will

admit the grease gun to the fitting without having to

remove the access panel.

Upper cowl panels installed on later serial aircraft

have an elongated slot in the access panel. This panel

P 465994-63 may

be

used to replace the earlier

access panel

P

465994-37. A A

~ ~ ~ ~ ~ ~ ~ ~

OMMER I L

SER V I E

ULLE T INS

PENDING

SUPER ONSTELL TION

SERIES

49

S8 No

2928

2976

2977

Approx

Release

May

1957

May

1957

May

1957

Subject

Replacement

of

Hydraulic

Pump Pressure Lines in

Nos.

1 2

and 4 Nacelles

Revisions to Curtiss Steel

Propeller Conduit

Revision to Vacuum Pump

Oil Separator Return Line

es ription of

Change

Similar to 1049/SB-2889 except is applicable only to

Nos.

2

and 4 nacelles see ig st Vol. 3

No.4 .

Provides new brush block and steel conduit fittings to

obtain adequate

pin

engagement see ig st Vol. 3

No.5 .

Increases the line size to preclude interchange with the

fuel vapor separator l ine see ig st Vol.

3

No.5 .

7

Documents

Lockheed Field Service Digest FSD Vol.3 No.6 Intro L1649 Starliner Part 2 of 3