was years ahead of any other country in the design and development
of intercity jet transport in 1950-51.
AVRO CANADA PLANT
August 10th 1949, the Avro C-102 jet transport, now better known
as the Jetliner, made its first flight.
aircraft was the first civil jet transport to fly on the North
American Continent and missed by only thirteen days the honour
of being the world's first jet transport to actually fly, which
went to the British De Havilland 'Comet'.
main purpose of this paper is to give a brief summary of the design
and general problems, which were encountered in the development
of the prototype up to the present flight test stage.
proceeding with the main portion of the paper, however, and with
apologies to the technical reader, the author would like to give
a short account of the events which immediately preceded the first
the 25th of July, the last stages of preparation for flight were
nearing completion and the aircraft had reached the stage of final
inspection and last minute test checking.
had worked on the project for almost three years and had of course
had numerous set backs, the biggest of which was the inability
of the engine manufacturers to suppoy the original twin engines,
and we had had to completely change the design of the aircraft
to accommodate four engines of a different type.
reached the final inspection stage, we thought therefore that
most of our bridges were crossed and all that we had to do was
get our aircraft into the air. It just shows how optimistic you
days later on July the 27th, it was announced over the radio that
the De Havilland "Comet" had made its first flight.
It was true that it had only hopped a few feet into the air, but
we realized that we had missed by just a few days the honour of
being the first people in the world with a true jet transport.
to make life still a little more complicated, the Department of
Transport deceided to tear up the runways at Malton, and carry
out extensive modifications, which were not scheduled to be finished
until some time towards the end of August. We were informed, however,
by DOT that we would have one runway on which to land, the 14-32
runway running north east and south west, with a bituminous surface,
and we would also have a short piece of concrete runway on which
to carry out our engine runs and taxi trails.
to confuse things still further, the tempurature deceided to take
a hand in the proceedings, and for several days before the first
flight was anticipated, it hovered between 90 and 100 degrees
engine runs having been completed over the week end, the aircraft
was wheeled out of the hanger on Monday Augest the 8th to start
taxi trails. As a special favor the tempurature had gone up to
103 deg F, nevertheless, we carried out our taxi runs, braking
tests, steering control tests, and towards early evening deceided
that it might be possible to attempt a hop, and take the aircraft
a few feet off the ground.
was not a very easy decision to make in view of the fact that
we had to contend with what was probably the highest temperature
of the whole year, and
with engines which were very much more susceptible to temperature
than normal reciprocating engines. We had a very short runway,
due to the alterations to the rest of the runways, and the pilot
was handling a completely new type of aircraft, the performance
of which could only be predicted at that time.
had calculated the distance required to take off and the decelerated
stop after the hop, and from our calculations, there were only
a few feet of runway left for pilot's error.
aircraft taxied down to the north east end of the runway, wasting
as little space as possible, the throttles were opened up, the
aircraft accelerated, and at about 90 mph, the nose wheel came
off the deck. A few seconds later, there were four loud and ominous
reports, the nose wheel came down, and the aircraft decelerated
to a stop, just a few feet from the far end of the runway. The
pilot had realized that he just could not make it and had applied
the brakes a little too early before the weight of the aircraft
was on the wheels, the wheels had locked, and all four tires had
blown out. In spite of this, the pilot had been easily able to
keep the aircraft on the runway, and there was no damage to the
wheels or brakes or any other portions of the aircraft.
aircraft was wheeled back into the hangar, the tires were changed,
and the next day more taxi runs were carried out to enable the
pilot to feel out the brakes before making another attempt at
Wednesday morning, August the 10th, three more runs were made
and a hop was attempted on the third run. This time, the two main
wheels on the starboard side of the aircraft blew out, and the
pilot again brought the aircraft to rest dead in the centre of
the runway, and this time with quite a bit of runway to spare.
tires were quickly changed and a conference held to decide whether
any more attempts at a hop would be made, as It was getting a
little expensive on the tires, and also on the nerves of the pilot,
co-pilot, and flight engineer who had to sit in the aircraft wondering
what was going to happen next.
pilot decided that the next time he went down the runway he would
rather take her up and "have done with it", as he expressed
it. The crew took time out for lunch, and after returning, decided
that in spite of a small gale that was blowing with quite a stiff
crosswind on the only available runway, and the fact that
the temperature was around
103degF, the next time they went down on the runway, they would
Just keep on going, and so just after lunch on Wednesday, August
the 10th, the Jetliner came down the runway, lifted off the deck
after a relatively short run, and gracefully climbed up to about
500 ft. where the pilot tried out the controls.
did a circuit of the field, and then asked for clearance to bring
her over the spot where the ground crew were standing to let the
boys have a look at the aircraft in the air. He then climbed
away to 8,000 ft. and reported after a few minutes flying,
that everything felt wonderful, and needless to say, everyone on
the ground felt pretty good too.
a flight of about one hour, during which time the aircraft was
flying at altitudes up to around 13,000 ft., the aircraft was
again seen, preparing for landing. By this time, the weather
man had turned on a crosswind of 35 mph at approximately 50deg
to the runway, but the pilot made an extremely short landing, and
taxied the aircraft down to the group of people waiting at the
was a general slapping of backs and congratulations all round,
the aircraft was wheeled back to the hangar and the first flight
of North America's first jet transport was
that time, the aircraft has done approximately thirty flights,
during which much valuable data has been accumulated, and the
aircraft is now well on the way to completing the tests, which
have to be carried out before the aircraft can be put up for C.A.A.
may be worth while mentioning one or two of the highlights of
this test programme.
most spectacular was probably the second flight, when after almost
an hour in the air, it was found that it was not possible to extend
the undercarriage, and it eas discovered later that this was due
to a fault in the main undercarriage gear. After losing most of
the hydraulic oil in the system, the pilot was forced to land
with the nose wheel down, the main gear up, and no flaps. The
fact that the flaps were up made the plane float, and the biggest
problem was getting it down at all, but after three runs, the
pilots brought the aircraft down on a grass verge at the end of
the runway, and skidded to a stop approximately 50 ft. from the
only damage sustained was four bent jet pipes and a caved-in plating
in the rear of the fuselage, and the landing only served to highlight
the safety of an aircraft which had no propellers to get in the
way on an emergancy such as this. The nacelles were repaired,
and the aircraft was flying again in just over four weeks, having
completed a test which no manufacturer would dare to carry out
at this stage in the life of a prototype, unless by accident,
as in this case.
is no doubt that we got a lot of data from this test, and we also
learned something from the tire bursting episode, as it was proved
that the aircraft could be brought to rest easily with any tires
burst, in any order.
series of tests probably worth mentioning are the engine cuts
at take-off. An outboard engine was shut down at various speeds
between 130 mph and 75 mph, it was still possible to take off
and have plenty of rudder power to spare.
lot of excellent data has also been accumulated on the low speed
charactistics of the aircragt. These have proved that the low
speed characteristics are just as good on a high speed aircraft,
if it is designed properly, as on the present conventional low
speed type of aircraft.
the jet transport had actually flown, there were many criticisms
of this type of aircraft, and some of them were so bitter that
one would almost think that they had been instigated by the manufactures
of propellers and their attendant controls.
of the criticisms was, that the runways and ground personnel would
probably get burned up when these aircraft were operating. It
would have done the critics good to see the official flight of
the Jetliner. On this day, the engines were started while the
aircraft was standing next to the big marquee containing the refreshments.
As the aircraft moved away, the people generally crowded in to
get a good look, and some of the press photographers appeared
to be almost trying to climb inside the jet pipe nozzles to photograph
the flames around the turbine, and nobody even got their eye lashes
point that has been grossly over-exaggerated is the takeoff and
landing distance required with the jet airliner. The Jetliner
has been repeatedly taken off and landed
at weights up to 57,000 lb. T. 0. gross weight during tests, in
distances of around 1,000 to 1,500 ft. and in one case, landed
at an average landing weight, less than 950 ft. from the approach
end of the runway.
tests on relightig procedures in the air have been carried out,
and engines have been shut down and restarted at various stages
during test flights, and it has never been necessary to attempt
more than one start on any engine. The results have been so good,
that it is now felt that relighting in the air is not only feasible,
but if carried out correctly is entirely without hazzard.
the most noticeable improvement inside the aircraft is the amazing
lack of noise. The test equipment for automatic recorder on the
Jetliner is about twenty feet aft of the cockpit, in the fuselage,
and the observer sits at this station with the various instruments
and cameras. It Is possible to converse without using the aircraft
inter-com by just carrying out a normal conversation between the
cockpit and the observers station.
one flight when a Lancaster aircraft was being flown along side
the Jetliner to get some photographs for the press,
the roar of the Merlin engines were quite apparent from
inside the Jetliner. It was almost possible to tell without looking
out of the window just how close the Lancaster was at any
lack of vibration is also very noticeable, and special vibrators
have had to be fitted on the instrument
panels to prevent instrument needles sticking. During the high
speed runs which were recently made at 30,000 ft., at which time
the aircraft reached speeds up to 500 mph, descent procedures
ware checked from 30,000 ft., and the aircraft was brought down
at a rate approximately 3,000 ft. a minute with the use of the
dive flaps fitted on the aircraft. There was no sensation of rapid
descent, and in fact, two of the observers had no idea that the
aircraft was descending at all, and were surprised to find themselves
at 20,000 ft. when they were at the opinion that they were taking
readings at 30,000 ft.
aircraft is at present being fitted with the necessary equipment
to test the air conditioning and pressurizing system, carry out
cruise control, and make a final assement of the aerodynamics.
To date, the test program has gone extremely well, and a large
ammount of data has been amassed in a relatively short time.
giving away any secrets, it can be said that up to the present
time, them has been surprisingly few snags, and to quote from
the pilots official report, "The aircraft has behaved magnificently
and is a very easy aircraft to fly".
following portion of the paper gives a brief history of the project
and covers some of the Technical problems encountered in the design.
the latter part of 1945, some interest was shown by the airlines
in both Canada and the United Kingdom in the remarkable
progress then being made with the use of the turbojet engine in
military aircraft. At the and of 1945. the Gloster Meteor was
in regular squadron service with the R. A. F., and the U. S. Army
Air Forces were also using jet fighters.
was generally agreed that if the advantages of high speed and
reduction of noise that the jet engine offered, could be combined
with the requisitesafety and economy essential in airline operation,
there would be a ready market for the high speed jet powered transport.
the spring of 1946, a detailed analysis was carried out at the
newly formed Avro Canada Organization at Malton, around a provisional
specification for a medium range inter-city turbojet transport.
The specification was based upon the requirements of the Canadian
domestic routes. The results of this analysis were sufficiently
favourable to convince both the airlines and the Company that
the idea of a medium range jet airliner was not only feasible,
it was also basically sound and should be proceeded with immediately.
design work was started in the summer of 1946 with an extremely
small design staff which was gradually increased, and by the early
part of 1947, the design was well under way.
order to reduce the number of untried features to a minimum, which
was obviously desirable both from the point of view of safety
and rapid development, the aircraft was designed on reasonably
was felt that the incorporation of too many design features which
had not been satisfactorily demonstrated on previous aircraft
would entail a considerable amount of laboratory testing, and
at the same time, the development costs involved would be prohibitive.
Nevertheless, enough original and novel design features were incorporated
to make the project unusually interesting.
the less conventional features will obviously be of the most
interest, these will be covered in greater detail in this paper.
general specification around which the aircraft was designed
was basically as follows:
|(1) The aircraft was
to be a turbojet powered short-to--medium range inter-city
transport with a still air range of at least 1,200 miles.
|(2) The payload was
to be at least 10,000 lb., and accommodation for not less
than 30 passengers was required.
|(3) A cruising speed
of over 400 mph at 30,000 ft. was specified without having
to resort to the use of oxygen for the passengers or
|(4) The aircraft was
to be designed to operate from airports with 4,000 ft. runways
under Standard Atmosphere conditions and comply with the
take-off conditions of the Civil Air Regulations. A decelerated
stop length of 5,000 ft. was not to be exceeded
under 'hot day' conditions following an engine failure.
at low speeds was not to be sacrificed in any way, despite
the high speed range required. The approach and stalling speeds
were to be at least comparable with present transport
|(6) Special attention
was to be given to serviceability and maintenance problems
to allow for maximum utilization and operational regularity.
|(7) The aerodynamic
and structural requirements of the Civil Air Regulations were
to be achieved.
|(8) The cost of operation was
to be comparable with or better than existing transports.
This then was the target. The figures in Table
1 will serve to show that it has not only
been achieved, but that the aircraft as now designed is superior
in all respects to the original specification.
T A B L E 1
C-102 JET TRANSPORT
|4 Derwent 5 Turbojet Engines
Total Static Thrust at Sea level
(I. C. A. N. Conditions)
| 14,400 lb.
|Gross Weight (Medium range version)
|| 65,000 lb.
|Gross Weight (Short range version)
|| 60,000 lb.
|Maximum landing weight
|| 52,500 lb.
|Still air range (Medium range
|| 2,000 miles
|Still air range (Short range
|| 1,400 miles
|Cruising speed at 30,000 ft.
and 60,000 lb. gross weight
|| 450 + mph
|| 12,700 lb.
|Number of passengers
|| 40 - 60
|Payload for 1,000
mile range,with full A. T. A. allowances
at 65,000 lb. T. 0. -gross weight
|Payload for 500 mile range with
full A. T. A. allowances
at 60,000 lb. T. 0. gross weight
|4 Engine take-off over 50 ft.
obstacle at 60,000 lb.
I. C. A. N. conditions sea level
|3 engine take-off with above
|Distance to Accelerate
to Critical Engine Failure Speed and Stop-ft.,
|(C. A. R. 04B.1221):
|60,000 lb. Gross, Weight at sea
|I. C. A. N. conditions
Distance from Height
of 50 ft.- ft.
|Sea level (I. C. A. N.)
|3,500 ft. (I. C. A. N.)
|Stalling speed at landing weight
of 50,000 lb.
with flaps in landing position.
|Stalling speed at landino, weight
of 40,000 lb.
with flaps in landing position.
To achieve the above results,
there were many difficult and new problems to be faced.
As there were no aircraft of this type in service, there
was obviously no experience or established data to fall
back on for many of these special problems.
A summary of some
of the major items, which had to be considered will
serve to show the nature of some of these problems.
obtain the optimum operating conditions with turbojet engines,
it is necessary to fly as high as possible. The reduction
in engine thrust between sea level, and say, 30,000 ft. is around
40%, while the drag is reduced to less than 25%, and as the thrust
from the engine is approxiamately constant for all speeds, the
variation being usually less than 5% between 200 and 500 mph,
it can be seen that flying at altitude is far more important than
with convential aircraft.
the interests of economy, it is also essential to climb the aircraft
to the operating altitude as fast as possible, and to descend
as rapidly as possible at the destination.
it would not be feasible to subject the passengers to the extremely
rapid changes of pressure caused by a quick descent, the pressure
in the cabin has to be constant as possible at all times. Statistics
indicate that average passangers when awake feel no discomfort
at equivalent rates of change in pressure up to 300 ft. per minute
in descent, and when asleep may suffer slight discomfort at a
rate of change of pressure somewhat below this. Most airlines,
therefore, recommend an equivalent rate of descent in terms of
pressure of not more than 200 to 300 ft. per minute.
conventional pressurized aircraft have the cabin pressurized to
8,000 ft. conditions at any altitude, 8,000 ft. being accepted
as the altitude to which the average person can climb without
feeling any discomfort either from lack of oxygen or reduced air
that this aircraft was pressurized to 8,000 ft. cabin conditions
at 30,000 ft., however, it would take 40 minutes for the aircraft
to descend at the recommended rate of 200 ft./min. This is obivously
not feasible with a jet aircraft, as not only would all the advantage
of speed be completely lost, but the fuel consumption of four
turbojet engines operating for most of the time at low altitude
would be prohibitive.
was obiously necessary, therefore, to pressurize the cabin to
as near sea level conditions as possible, right up to the cruising
altitude to enable the aircraft to be brought down in the shortest
possible time. The conditions achieved to date are as follows:
a sea level cabin up to 21,250 ft., a 2,000 ft. cabin at 25,000
ft., and a 4,000 ft. cabin at 30,000 ft. The pressure differential
to achieve this is 8.3 lb./sq.in., and as a safety factor of 2
is used for pressurizing, the fuselage had to be designed to withstand
a pressure of 16.6 lb./sq. in. The structural problems involved
with the use of these high pressures were to say the least, interesting.
it is obiously not desirable to put large access holes and doors
in the fuselage for servicing under these pressures, a great deal
of ingenuity had to be used to cut down the number of external
holes, and at the same time design for efficient servicing, and
decompression due to a window blow-out etc., is always a problem
in considering high altitude flying for passanger carrying aircraft.
It is comforting to note, however, that in the opinion of the
Aviation Medicine experts, the only real physiological discomfort
up to 30,000 ft. is the lack of oxygen. Above 40,000 ft., the
average individual is unable to obtain sufficient oxygen, even
when breathing an atmosphere which consists entirely of oxygen,
because of the decrease in total pressure in the lungs.
the optimum operating altitude of the C-102 was set at 30,000
ft. bearing in mind the best flight path for average range, this
problem was not considered to be too serious. Investigation is,
however, going ahead on the basis of an automatic oxygen system
which comes into operation, if a blow-out does occur, and which
floods the cabin with oxygen vapour.
CHOICE OF ENGINES
designed as a twin engined transport, the C-102 was designed to
take two Rolls-Royce Avon engines.
the fall of 1947, when it was realized that the Avon engines would
not be available for the first prototype, it was decided that
four Rolls-Royce Derwent engines would be used on the first aircraft. The
decision to do this was not taken lightly, as it involved a complete
redesign of the centre section structure which was then somewhere
near design completion. The sideways retracting undercarriage
scheme had also to be completely, scrapped.
was necessary to start from scratch on the nacelles, and the change
in centre of gravity due to the addition of two extra engines
necessitated a repositioning of the wing in relation to the fuselage.
the redesign work progressed, however, it became evident
that the use of four engines was not only a very much better and
safer arrangement, but the fact that the undercarriage would now
be retracted fore and aft in the nacelles made the undercarriage
unit and adjacent structure very much simpler in all respects.
Also the use of engines which had been operating in military.
aircraft for over 100,000 operational hours was a very big point
in eliminating one of the big unknowns, which would have had to
be faced with the use of engines which were only in the development
use of four engines also made compliance with existing C. A. A.
requirements very much easier, and the engine failure case less
severe on the control surfaces.
decision to use an underslung nacelle instead of the "through-the-
spar" arrangement necessary with the original engines, also
simplifies the fitting of newer types of engines as they become
available without any major structural alteration.
high speed and relatively low wing loading resulted in the load
factors being considerably higher than those at present used for
transport aircraft. Reference to C. A. A. 04. 21411 shows that
the gust factors vary directly with the speed and inversely with
the wing loading,
relatively large amount of fuel carried in the jet powered transport
resulting in a low landing weight, and consequently, a low wing
loading, together with the increased speed, all make for a higher
limit load factors for gust conditions can be seen In figure 2,
and these have to be multiplied by a safety factor of 1.5.
highest limit load factor is 4.5 at an empty weight of 34,000
lb. and a speed of 300 mph E. A. S.
The overall wing loads were also increased due to
the absence of relieving loads from conventional outboard nacelles.
compensate for the increased structural strength required, the high
strength aluminum alloys 75ST and 24ST are used extensively to obtain
the maximum strength-to-weight ratio.
outer wings are also designed as fully stressed skin structures
with heavy gauge skin and stringers taking the place of the usual
concentrated spar booms, and providing a high depree of torsional
heavy skins are used on the lower portion of the fin to give torsional
rigidity and prevent tail flutter.
windscreen structure is a high strength aluminum alloy casting,
and the pressure bulkheads are situated at the front of the windscreen
and at the rear of the passenger cabin.
rest of the structure is along conventional lines, and so will not
be dealt with in any great detail. The structure weight is approximtely
27% of gross at 60,000 lb.
GENERAL AERODYNAMIC CONSIDERATIONS
The reduction of the parasitic portion of the total drag is most
important with the turbojet powered aircraft.
Reference to figure 3 will show that the ratio of fuel consumption
to thrust does not increase very rapidly for speeds between 300
and 500 mph.
the thrust is approximately constant for all speeds, it is apparent
that the miles travelled per unit of fuel is increased in almost
direct ratio to the aircraft speed. The aircraft then, has to be
aerodynamically clean to cut the parasitic drag to a minimum.
the design of the C-102, the greatest care has been taken to get
a good external finish, and all external riveting is flush. The
skins are pre- formed and stretched to provide the smoothest contour
and practically all the radio antennae are flushed into the contour.
The exceptions are the short radio compass sense antenna in the
nose, and the conventional wire antenna for H. F. communication.
choice of wing section is always of necessity a compromise and the peculiar
conditions which had to be met, with a transport which was
to be almost twice as fast as existing transports made this problem
even a little more complex than usual.
was obviously essential to cut down the drag to a minimum, and at the
same time to obtain the highest possible CLmax for take-off and
structural problems with high gust factors and the large amount of
fuel which had to be carried also influenced the wing design.
section chosen to obtain the best all round characteristics was
a relatively high cambered aerofoil, with a thickness at the root
of 16.5% and 12% at the tip. The aircraft will be operating at a
mach number of less than .7 at 30,000 ft., and no compressibility
problems are expected with this aerofoil at these speeds.
aerofoil also has the advantage that the trailing edge angle is
low, and the pressure recovery gradient is conservative, which makes
the section less sensitive to manufacturing and junction interference.
Wing Plan Shape
fairly low wing loading was used for better approach characteristics,
and the plan shape which appeared to give the best compromise was
one with an aspect of 8.1 and a taper ratio of .5. As the basic
characteristics having the greatest effect on stalling is the taper
ratio, this was chosen very carefully. The straight center section
makes the fuselage-to-wing junction easier to manufacture and helps
in the power plant installation as the engines are on the parallel
was considered that for an aircraft operating at a Mach number less
than .7, sweep back would not be worth the extra weight which it
would envolve. The best arrangement appeared to be that having a
straight rear spar, which gives a sweep back of approximately 4
1/2 deg at the quarter chord. Washout was considered, but did not
seem to give any great promise, as although it gave slightly better
stall characteristics, the effect of the extra induced drag at high
speed was less favourable, and the manufacturing difficulties would
also be very much greater.
flaps are fitted on the first set of wings for the first prototype.
These will later be changed to the double-slotted type to cut down
landing, and approach speeds to a minimum. As these are of the area
increasing type, there will be a larger center of pressure movement
than with the split flaps, and slightly greater elevator angles
to trim may be required when these are fitted.
profile drag has been kept to a minimum by the use of thick skins
required for wing stiffness, and complete flush riveting. Square
tips are used, to give greater aileron effectiveness by carrying
the surfaces out as far as possible, and for ease of manufacture
of the tips themselves.
dihedral on the outer plane is 6 deg., and the wing incidence is
2 1/2 deg., throughout the span.
shape of the fuselage is the usual compromise between getting a
profile which is aerodynamically clean and a structure which is
easy to assemble, coupled with the standardization of interior fittings
and structure for as long a lenght as possible. This has resulted
in a parallel section of fuselage for approxiametly 60% of the total
lenght with a carefully blended-in fore and after-body.
care was taken to get good lines around the nose canopy, and wind
tunnel results showed that the critical Mach number around the canopy
is about .73 i.e., higher than that for the wing.
00-21 series aerofoil section has been used for the after-body,
which again showed very good pressure recovery characteristics in
the wind tunnel. A circular cross-section is used throughout, as
this is obviously the best section for the high pressure differential
used on this aircraft.
shape of the Dorsal fin was determined by the requirements for weather
cock stability and helps to co-opt a portion of the rear fuselage
as fin area.
Tailplane Vertical Position
tailplane is located high on the fin. If the tailplane was on the
center line of the fuselage, it would be directly in the wake of
the jets. While the temperature effects of the jet stream are not
too serious by the time they get to the tail, the velocity effects
are more marked. If the tail was just out of the jet stream, but
fairly low down on the fin just above the fuselage, there would
be a marked interference between the sharply tapered after-body
and the tail-plane.
Effect of Thrust on Trim
jet nozzles are inclined at an angle of 7 deg to bring the line
of action of the jets as close to the normal C.G. position as possible,
and minimize the effect of change of trim between power-on and power-off.
jet stream has a cleaning up effect around the trailing edge of
the center section wing. When the engines are opened up during a
baulked landing, air is drawn into the jet stream over the adjacent
wing surfaces due to the greatly increased velocity through the
jet nozzles. This has the effect of reducing the stalling speed
under these conditions.
Wing Root Fillet
unusual design of wing root fillet eas incorporated to take care
of the upwash from the fuselage. The
normal component of the flow around a long nosed fuselage produces
an upflow at the wing root, which may cause premature root stalling,
and during wind tunnel tests, it was found that a long forward fillet
of the right shape corrected this effect, see figure 14.
fillet was tried out on a British aircraft by arrangement with Avro
Canada and produced excellent results. The stalling speed was reduced
by approximately 7 mph E.A.S. after incorporation of the fillet.
There was no effect on the longitudinal stability.
the various flight plan proceedures have been worked out between
the airlines, the Civil Air Authorities, and airport control personal,
it is obiously not possible to give a definite flight plan. Figure
4 shows a recommended proceedure, which allows for a standard 45
minute stacking and 120 mile flight to an alternative airport, plus
an allowance for instrument approach, landing and taxiing.
will be seen that instead of the usual proceedure of descending
at the destination airport, taking a pass at the airport to check
whether the landing is possible, and then proceeding to an alternative
airport, the decision to descend or proceed to the alternative is
made at some point during descent.
point is shown on the flight plan at an altitude of 25,000 ft. and
approximtely 33 miles from the airport, and this is considered to
be entirely reasonable with present ground aids and radio equipment.
to the high cruising speed, it is expected that the weather at the
destination will have been reasonably accurately established, and
will not have changed during the short flying time.
conditions are considered to be unfavourable for landing at the
destination airport, the aircraft proceeds at its best endurance
speed at an altitude of 25,000 ft.
stacking required is carried out at an altitude of 25,000 ft., or
could be carried out at any altitude on two engines, without any
penalty in fuel consumption. When the aircraft is given the signal
to land, the normal procedure of descent and instrument approach
is then made at the alternative.
flight plan as shown is applicable for all ranges above approximately
200 miles. For ranges under 200 miles, It is debatable whether it
is worth while climbing to an altitude of 30,000 ft. for cruise.
will be seen that for a range of 500 miles, the take-off and climb
to 30,000 ft. covers a distance of 90
ground miles, and the descent from 25,000 ft., approximately 33
miles. Normal cruise at the operating altitude covers approximately
fuel used for take-off, climb to 30,000 ft., cruise, descent and
approach for a range of 500 miles is approximately 9,210 pounds,
while the fuel allowances carried for flight to alternative, stacking,
and descent at alternative airport amount to approximately 5,125
pounds, or just over 1/3 of the total fuel.
is carried out at a speed of 200 miles an hour E. A. 8. with the
use of dive brakes to get a high rate of descent.
the accessories, including the hydraulic pumps and electrical equipment
for de-icing may be required during the descent, the engines are
throttled down to 7,000 r.p.m., at which speed, the accessories
are designed to maintain the full output required for any of the
is very little penalty on rate of deseent incurred by keeping the
engines running at this r.p.m. Figure 3 shows the rate of descent,
power-off compared with cruise r.p.a. and 7,000 r.p.m.
shown, there is little difference between the power-off and half
max. cruise engine speed curves. At this speed, the cabin blowers
will also give their full ventilating output, and in any case, will
be operating at a rapidly reducing back pressure during descent.
average fuselage angle during descent is not more than 8 deg, which is
considered to be reasonable from a passenger comfort standpoint.
DIRECT OPERATING COSTS
it is not the purpose of this paper to join in the merry-go-round
of comparisons of the conventional and jet-powered transports on
a ton-mile per lb. of fuel basis, nevertheless, the operating costs
had to be considered very carefully, and their consideration played
an important part in the final design confimration of the aircraft.
two important efficiency factors in the cost analysis are the cost
per mile and the payload for a given range.
cost per mile is obviously governed by speed, as manv of the direct
costs such as, crew salaries, depreciation, insurance, etc. are
fixed hourly costs. Neglecting fuel consumption, if the blockspeed
is increased from say 250 mph to 350 mph, the cost per mille
would be decreased by approximately 30%. It can and has been
shown elsewhere, that this decrease in cost due to speed more than
compensates for the increase due to higher fuel consumption.
effect of blockspeed can possibly be seen more clearly by considering
the number of aircraft required for a given scheduling. The equation
in its simple form is shown below.
U x Vb x Np
|Number of aircraft required
|Traffic density in passenger miles
|Utilization in hours per year
|Passenger capacity of aircraft
a given yearly utilization, traffic density and passenger capacity,
it can be seen that if the blockspeed is doubled, the number of
aircraft required is halved, and consequently, the earning power
of each aircraft is considerably increased.
take advantage of the higher blockspeeds, however, maintenance and
turn-around time at the airport has to be cut down to a minimum,
and the optimum climb and descent procedure from operating
altitude taken into account.
high degree of pressurization and the incorporation of dive
flaps to allow a rapid descent; the use of special accessories and
radio compartments where practically all items that required frequent
servicing are housed; and the employment of underwing pressure refueling
are only a few of the items, which have been incorporated to increase
the economic efficiency of the aircraft.
far as the payload portion of the cost per ton mile efficiency datum
goes, the fuselage was laid out to give the best compromise
between a full passenger version and combined passengers and cargo.
Two typical lavouts are the 40 passenger version with an additional
4,100 lb. of freight making a total pay-load of 12,500 lb., and
the 50 passenger version with a payload of 10,500 lb. Payload Vs.
range with all ,allowances is shown on figure 6.
the final analysis of economy must be left to the individual airline,
the results of a detailed analysis show that the direct operating
costs compare very favourably with those of present transports,
despite the relatively high fuel consumption of present jet engines,
and the fact that the present allowances for stooge and flight to
an alternative airport are severe on the jet transport.
is obvious that as the specific fuel consumption for the jet engine
improves, with the use of ceramic blade materials, and higher compression
ratios, and the flight procedures are modified to cut down the stooge
time, the picture will be even brighter.
the final seating arrangement and cabin layout will depend on the
customer's choice, it appears to be fairly definite that the high
density passenger version will be the one of greatest interest,.
typical layouts are shown in figure 7. Accommodation for 40 or 50
passengers is shown with provision for their baggage on the left
hand side of the cabin, adjacent to the front entrance door. The
situated opposite this baggage compartment, and a small commissary
and hostess station is situated at the rear of the passenger cabin.
ten ft. diameter fuselage allows for wide seats, and a generous
aisle with a head room of 82 inches. Seat pitching is at 38 inches.
exits are situated in the centre section and rear section above
the wing, and a crew emergency exit is fitted in the ceiling of
the crew compartment.
permissible C. G. travel of approximately 23%.of the mean chord
or approximately 35 inches provides for flexibility of loading with
the least number of restrictions.
Noise Level and Vibration
level in the cabin is considerably reduced by the use of turbojets,
and this, coupled by a complete lack of vibration, will add enormously
to passenger comfort.
four Derwent 5 engines are mounted in pairs in two under-slung nacelles,
each nacelle being made up as a single integrated structure. The
engines are toed-in toward the centre line by 5 deg, and set at
approximately 11 1/2 deg to the horizontal in order to take the
jet pipes under the main spars without cutting away
any of the spar structure.
engine mounts are used, and these can be removed or replaced separately.
The nacelle geometry is shown in figure 8. All nacelle air loads
are taken back into the two engine mounts, which are attachedto
the centre section front spar.
servicing and maintenance is made particularly easy by the low position
of the nacelles. All engine
adjustments can be made without the necessity of using service ramps
or ladders, see figure 9.
removal is carried out by detaching the services and gear drive
at the break points, swinging the trunnion locating caps down, and
dropping the engine on to the special trolley. The engine is then
wheeled away sideways to make way for the replacement engine. With
this unique arrangement, a complete engine change can be made in
a very short time.
jet pipes are parallel in plan and are supported on trunnions and
links. Two spherical joints are incorporated to give flexibility
to the pipes on expansion, and also for the withdrawal of the jet
pipe for engine removal. Although, a relatively long jet pipe is
used, it Is estimated that less than 1% of thrust is sacrificed
from the combined effects of length and shape of the pipe.
sixteen inch nozzle is fitted and the jet emerges at 7 deg. to the
datum line of the aircraft, to bring the line of action of thrust
as close to the C. G. as possible.
jet pipe runs through a tunnel of stainless steel formed by firewalls
attached to the adjacent structure. The jet pipe itself is insulated,
and is cooled by a flow of air passing through the firewall tunnell
and induced by the extractor nozzle. The vena-contractor at the
nozzle sucks the cooling air through the nacelle, after it enters
through louvres at the forward end of the cowling.
main accessories driven by the engines are mounted on an accessory
gearbox located between the engines in each nacelle,and attached
to the wing front spar. The gearbox contains two completely independent
gearing systems, each driven by one engine, and each having independent
inboard engine drives a cabin blower, a vacuum pump, and a Tachometer
generator, and each outboard engine drives a 50 KW alternator, a
9 KW generator, an hydraulic pump, and a Tachometer generator.
gearbox drives are connected with the engines by a system of drive
shafts linked by means of flexible couplings, as shown in figure
Derwant 5 Modified Engines
C-102 engines are standard Rolls-Royee Derwent 5 engines but with
a completely redesigned oil tank.
The cast oil tank is sited at the front of the engine, underneath
the forward gear drive. The engines are handed only by the oil tank
filler and the gear take-off, see figure 11.
change from starboard to port engines is made simply by interchanging
the filler neck and blanking plate on the oil tank and swinging
the gear take-off around in the opposite direction. The oil tank
and system are integral parts of the engine.
engine is supported by mounting trunnions at approximately the center
of gravity of the engine, and is steadied at the rear and by a shackle
plate bolted to the top of the nozzle box.
upper part of the cowling is developed as a permanent structure
provided with small access doors for engine slinging, and a larger
door to permit access to the upper part of the accessory gearbox.
lower half of the cowling consists of two large access doors hinged at
the sides and a smaller door beneath the accessory gearbox swinging
aft. All access doors are locked by means of flush-type quick release
fasteners, and the two main curved doors can be quickly
detached by swinging them out and unhooking the special
the engines installed, the nacelle is divided into two compartments
on each side, and a third compartment rousing the accessory gearbox.
This split-up is achieved by means of special, firewalls and bulkheads
as shown in figure 8. Each nacelle has a vertical firewall forming
a centre keel and isolating the two engines from each other.
engine has an integral intermediate firewall permanently attached
and sited around the combustion chambers. This mates up with a permanent
portion of firewall on the nacelle foming a complete firewall between
the hot and cool portions of the engine.
front portion, or zone 1, which also forms the plenum chamber contains
the engine accessories and oil tank etc., while the rear portion
or zone 2, contains all the hot portions of the engine, combustion
chambers, turbine casing, and jet pipe. The intermediate firewall
is to prevent the high pressure fuel from a burst pipe or joint
being sprayed on to the hot side.
rear portion of zone 2, extends in the shape of a tunnel back to
the jet nozzle and is completely lined with stainless steel firewalling
and sealed against ingress of fuel or oil.
from a burst combustion chamber or perforated jet pipe would be
confined within this zone out of reach of electrical and fuel lines
or the aircraft structure.
above system of firewalling also isolates all engine parts from
the accessories and gearbox, which are in the space above the conical
firewalls, shown on figure 8, as zone 3.
resetting type fire-detectors are used and a methylbromide system
of extinguishing is used for zones 1 and 2, while a C02 system is
provided for the gearbox compartment, zone 3. A two-shot system
is used and the warning lights, buttons, and selector switches are
mounted on the ceiling fire-protection panel in the cockpit.
thrust from a jet engine varies considerably with temperature and
airport altitude, and on a hot day with
a temparature of 110 deg F, the reduction in jet thrust can be as
much as 16%. As this can be critical for take-off conditions, where
a possible engine failure has to be taken into account, some means
of thrust augmentation has to be used.
means of achieving the extra thrust were investigated, and it was
finally decided that injection of a water-methanol mixture into
the compressor inlet offered the best solution. The predominant
effect of this is to increase the mass flow of air to the engine
by increasing the air density at the compressor inlet.
injection system itself is relatively simple, and has few of the
disadvantages of other forms of augmentation such as, after-burning
where the long sheets of flame coming out of the tailcone are likely
to cause alarm to the passengers. The percentage increase in thrust
with rate of injection is shown in figure 12.
can be seen from the graph, that under tropical conditions, to provide
the static thrust which would be obtained for take-off under standard
I. C. A. N. conditions, it is necessary to inject the mixture at
a rate of 10
gals. per minute.
tank is housed in each nacelle holding 66 gals. of water-methanol,
which is sufficient to supply each engine vdth the required quantity
for a period of three minutes.
13 shows the take-off distance for various gross weights and temperatures.
external and internal shape of the nacelles was chosen very carefully
with a view to getting the best possible pressure recovery characteristics
externally, and an efficient plenum intake which would give the
best compromise between the ideal low and high speed conditions,
see figure 8.
take-off conditions where there is very little ram effect, there
is a suction
in the plenum chamber, and in order to prevent breakaway around
the intake walls, the wall angle was kept down to less than 10 deg.
To achieve this, it was necessary to go to separate, intakes foreach
engine, as with a common elliptical intake, the diffusion angle
would have been excessive in a short nacelle, and any increase in
nacelle length was disadvantageous, due to the destabilizing effects
of a long wide nacelle.
best intake curves were established in conjunction with the engine
manufacturers recommendations. For the outside shape, the lines
between the inside lip of the intake radius and a point about 20%
of the total nacelle length aft of the intakes were most critical
both for drag rise and intake effeciency, see figure 14. Figure
15 shows how little the nacelles interfere with the top surface
of the wing.
civil version of the standard Rolls-Royce Derwent 5 engine is used,
and a brief summary of the performance is shown below.
|Take-Off and Climb
|Maximum continuous power
|Idling on ground
in the air is possible and numerous relights have been carried out
during flight tests.
the economy of the C-102 has been worked out assuming that all engines
are operating, however, relighting would not normally be employed.
It can be seen by reference to figure 16, that each engine consumes
less than 90 lb. of fuel in descent from 30,000 ft. at half max.
cruise r. p. m. If the operator felt, however, that any stacking
should be carried out at fairly low altitude, two engines could
be closed down to conserve fuel.
is housed in four integral wing tanks located in the inboard portion
of the outer wings, between the main spars. The total capacity of
the tanks on the first prototype is 2,400 Imp. gals. The tank capacity
can, however, be considerably increased. Immersed booster pumps
pilot can fully control the disposition of his fuel load, and a
cross balance pipe is provided so that fuel from any tank is available
to all engines in an emergency.
the event of failure of the booster pumps, the engines are capable
of sufficient suction to enable them to operate with the booster
controlled shut-off cocks to each engine are provided as a safety
measure to shut off the fuel in the event of an emergency.
signal light system is provided on the, fuel system panel to enable
the pilot to check instantaneously the condition of the fuel system.
overwing and underwing refueling is installed and the tanks can
be refilled at the rate of 200 Imp. gals. per min. through each
underwing refueling valve at a nozzle orifice pressure of approximately
5 p. s. i. A refueling manifold is used for each pair of tanks and
a special built-in selector valve permits fueling or defueling of
each tank individually.
special float valve coupled with the underwing refueling system
prevents the tank being damaged, by shutting off the fueling valve
when the fuel reached a predetermined level in the tank.
scaling of the integral tanks was a problem which had to be studied very
carefully, since the airlines had been having some trouble with
certain types of integral tanks and a certain amount of prejudice
had been built up against them.
much investigation and testing, a system of sealing was derived which
has given such excellent results on test that it appears to be a
very great improvement on the existing methods of sealing.
sealants are used and combinations of plasticizers and synthetic
resins are added, making
a permanently plastic seal, which has low shrinkage and good
adhesion properties. The top and bottom wing skins and the spars
are sealed before assembly, and the corners are then sealed after
the wing is removed from the assembly jig.
sealant is used between the faying surfaces. The finished tank is
sprayed with a cyclohexanone solvent to bond the complete inner
surface. The system lands itself to local repair as no slushing
compounds are used
to the wing is by large leading edge access doors and stress-bearing
removable panels in the front spar, see figure 17.
aerodynamically-unbalanced control surfaces have been used for both
the rudder and elevator controls, see figure 18.
The intermediate or auxiliary surface on the rudders
is used soley to trim out for an engine failure at low speeds. With
the use of jet engines, high rudder angles are not normally necessary
due to the absence of slip stream,
which is the usual cause of swing at take-off. The engines are also
close to the fuselage which again reduces the rudder power required.
tail plane is out of the flap wake during landing and, therefore,
the tail efficiency is high which reduces the elevator angles required
for normal trim. The auxiliary surface is only required for the
flare-out, on landing with an extremely forward C. G. Piano hinges
have been used on all tail surfaces, and this improves the effectiveness
by sealing the gaps.
chord high aspect ratio surfaces are used, and these have the advantage
that no aerodynamic balance is necessary. They also have lower drag,
less danger of icing, better repeatability and low weight of mass
narrow chord elevator is also very much better from the point of
view of susceptability to oscillatory instability. The usual cures
for this are less aerodynamic blance, and a lower mass moment of
inertia. These features are all incorporated in the double surface
operation of the tail surfaces on the first prototype is by a simple
switch controlling a small electric motor and limit switches. The
system is entirely separate from the electric and manual elevator
hydraulic assister is used for aileron power boost in the ratio
of 5 to 1. This is a pure assist system, and in the event of an
hydraulic or unit failure, the booster is thrown out and full manual
control is retained with, of course, reduced power.
type controls are used on all three main control systems, employing
light alloy tube to eliminate differential expansion and contraction
under extreme temperature changes. The tubes are supported in roller
guide bearings using rubber covered ball bearing rollers.
air conditioning system is entirely automatic once the controls
have been pre set by the pilot. Either supercharger is capable of
delivering about 60 pounds of air per minute up to an altitude of
13,500 ft. Automatic control of the cabin pressure is maintained
by the discharge valve set to provide sea level conditions up to
21,500 ft. the differential pressure remains constant, and at 25,000
ft., the cabin altitude is 2000 ft. and 4,000 ft. at 30,000
rate of pressure change in the cabin during the climb and descent
is also automatically controlled.
fuselage had to be very carefully sealed to provide a pressure tight
cabin and a method of sealing
was used which has been well tried on other aircraft.
consisted of applying special sealing compounds between the faying
surfaces and skin joints. The remaining riveting such as, riveting
stringers and capping strips to the skin were not sealed, as with
the use of dimpled riveting, the rivets are tight enough to produce
a satisfactory seal. Any leaking rivets are individually sealed
by bushing with a special sealant.
19 shows the cabin insulation installed prior to fitting the wall
in mind the usual confusing array and disposition of instruments
and controls in the average flight deck, a special attempt was made
in the case of the C-102 to achieve a configuration that was both
functionally good, and at the same time, gave the best servicing
extent to which this has been achieved can be seen in figure 20.
The main instrument panel is divided into three sections. The centre
panel carries all engine and fuel instruments. A small fuel system
control panel is attached to the engine panel with the fuel diagram
etched on, and this contains the switches and lights for
the various booster pumps and cross-feed warning lights. All panels
are hinged for easy access.
engine instrument panel is very much simplified by the use of jet
engines, as the only engine instruments are the R. P. M. indicators,
jet pipe temperature gauges, burner pressure gauges, and oil pressure
two main instrument panels carry the normal flight instruments,
and have been grouped to conform with the latest requirements for
radio navigation and automatic landing aids.
the ceiling, between and within easy reach of each pilot, is the
main electrical panel carrying the engine starter switches, fire
protection switches and buttons, and the main electrical control
pressurization control panel is on the left of the captain and the air
conditioning, oxygen and de-icing control panels to the right of
the first officer. Circuit breaker panels for both electrical and
radio equipment are mounted on the aft deck bulkhead.
pilots' seats are fully adjustable and slide back for easy access.
Cranked control columns are used to avoid obstruction to the pilots'
knees, and a spectacle type of aileron hand wheel is used.
lot of thought was put into the main control pedestal, which on
the upper portion carries the engine throttles, undercarriage, flap,
and automatic pilot controls, the emergency manual low pressure
fuel cock levers, and fuel tank selectors.
radio control panels are situated on the lower portion of the pedestal.
The pedestal also carries all the manual trimmer controls, the manual
autopilot disconnect lever, gust lock and parking brake levers,
and the aileron power boost cut-out.
vision windows which swing inwards are provided for landing under
adverse weather conditions.
rudder pedals are fully adjustable and are articulated to provide two
brakes for equal or differential brake application.
above cockpit layout was finalized only after many conferences with
airline pilots and technicians and the final mock-up was carefully
checked to get the best possible layout.
absence of propellers and the consequent short distance between
the aircraft structure and the ground coupled with the fact that
an under-slung nacelle configuration was used, resulted in
an extremely short main
landing gear, see figure 21. The actual distance between the undercarriage
main pivot and wheel centres is less than 30 inches. This has resulted
in the establishment of an extremely robust and simple design, and
one which is believed to be lighter as a percentage of the gross
weight than any existing transport undercarriage.
wheels are used on both the main and nose units, both retracting
forward. The main undercarriage struts consist of a telescopic leg
incorporating liquid springing. The nose wheel is self-centering,
fully castoring and incorporates shiming damping, see figure 22.
The hydraulic steering unit incorporating a double piston control
acts as a shimming damper when the steering is switched off. Steering
is controlled by a wheel adjacent to the pilot.
nose wheel is steerable through an are of 70 deg each side, and
the wheel cancaster through 360
deg for towing. Lever suspension is used on the nose gear.
retraction is electro-hydraulic and hydraulic brakes have been installed,
controlled by the rudder pedal toe brakes. All wheels are to the
American Tire and Rim Association specifications. The main undercarriage
doors are operated by separate hydraulic jacks, and the nose wheel
doors are operated mechanically by a trip mechanism fitted to the
ground retraction is prevented by a microswitch which comes into
operation when more than 5% of the aircraft weight is on the wheels.
undercarriage uplocks can be tripped manually in an emergency and
extension will then take place by gravity and drag forces.
position and layout of the various accessory units which have to
be serviced regularly on the ground,
or which need to be accessible in flight was given a lot of thought,
as this is a point, which has aroused much criticism in the past
by airline operators.
accessories compartment was introduced behind the first officers
bulkhead on the starboard side to carry the main aircraft accessories,
see figure 23. The heater, refrigerating turbine, main electrical
accessories such as, inverters, relays etc., and the main electrical
distribution panel are all housed in this compartment, which has
its own fire extinguishing system.
radio and electronic units are in a similar separate compartment
on the port side behind the pilot's bulkhead,
see figure 24.
main hydraulic units are panelized, the panels being housed in the
forward wing root fillet, with easy access at ground height to all
ground connections, accumulators, valves etc. The emergency power
pack is also contained on these panels.
engine fire protection bottles are housed in the nacelles at shoulder
height and the engine starter relay panels are also in this vicinity.
extremely low static position of the aircraft imures that
practically all external servicing is done without steps or servicing
pressure tests can also be carried out by connecting up the ground
pressurizing equipment to a service panel inside the nose wheel
main hydraulic system is a high pressure system operating at a normal
pressure of 1,800 pounds per sq. inch. The cut-out pressure is 2,200
p. s. i. and the relief valve pressure is 2,700 p. s. i.
normal system power is provided by two constant pressure variable,
displacement pumps on the accessory gearbox in the nacelles. Either
pumps will provide full hydraulic power for the complete system,
and the use of two pumps is to provide duplication against failure.
main services operated by the hydraulic system are the main and
nose undercarriage gear, nose wheel steering unit, landing and dive
flaps, main wheel brakes, main wheel doors, and aileron power booster.
Complete duplication of the normal hydraulic system is provided
by a "power pack" consisting of an electrical motor and
hand pump is provided in the accessories compartment which also
can be used in an emergency. On the ground, the system can be operated
by ground supply points located on the wing root fillet panels.
electrical system is basically a single grounded negative system
for both D. C. and A. C. services.
are in effect six separate systems providing power for the various
services. The various systems are listed below:
volts - From two engine driven D. C. generators for lighting
relay controls, radio and some instrumentation. The D. C. generator
system is over-voltage protected.
115 volts - Three phase 400 cycles from D. C. motor generators
(inverters), for some flight instruments, engine instruments
and some radio equipment.
volts - Three-phase 400 cycles from a transformer, connected
across the 115 volt three-phase power supply for
volts - Three-phase 400 to 700 cycles from two engine driven
alternators for wing and ampennage de-icing and galley.
volts - Three-phase 400 to 700 cycles, from a transformer
connected across the 208 volt three-phase power supply for the
'Nesa" de-icing system.
89 ampere hour batteries connected in series to supply 24 volts are
used for ground testing and generator stabilization in flight.
de-icing will not be fitted for the first flights of the first prototype,
an electro-thermal de-icing system will be used for the wings and
empennage. De-icing power is provided by two 50 KW., 208 volt three-phase
400-700 cycle alternators situated on the engine driven
Windscreen de-icing is provided bv special 'Nesa'
glass windscreen panels, which consist of a vinyl core sandwiched
by two thicknesses of semi-tempered glass. On the outside surface
of the vinyl between the vinyl and the outside layer of glass is
a conductive 'Nesa' coating which provides approximately 5-6 watts
per sq. inch power input.
windscreen de-icing is entirely automatic and the temperature is
controlled to provide the quantity of heat required for anti-icing,
and at the same time, keeping the vinyl layer at a temperature which
gives it the best resistance to bird impact.
three forward panes of the aircraft are designed in this manner,
and the vinyl centre layer has the additional advantage, that in
the event of a windscreen being shattered by any circumstances,
the vinyl will still withstand at least twice the maximum differential
pressure in the fuselage by blowing out in the form of a bubble.
and intake de-icing is a special problem which at the moment is
being investigated fully by the engine manufacturers and Avro Canada.
There are several workable schemes. As it has not yet been decided
definitely which system will be used for production aircraft, it
is obviously not desirable to go into detail on the subject in this
radio equipment is housed in the radio compartment on the starboard
side of the front entrance door, and is completely enclosed by quick
removable panels giving complete access to all units.
electronic units are housed on sliding racks and use is made of
special type connector boxes which automatically engage the pins
when the units are pushed into position.
radio and electronic compartment is ventilated by a separate blower.
The basic radio system consists of the following:
|(1) HF communication transmitter-receiver with
provision for 20 channel equipment.
|(2) VH communication transmitter-receiver.
|(3) 18 channels plus guard channels.
|(4) Dual automatic radio compasses with radio
|(5) Isolation amplifier chassis including interphone
amplifier,and a special loud speaker amplifier for the
captain and the first officer.
the radio and navigation instruments are duplicated on the captain's
and first officer's panels, and all control panels are located on
the flight deck pedestal.
entire radio pedestal assembly is removable as a unit by means of
disconnecting plugs at floor level.
and headphone jacks are provided at the side panels, and separate
loud speakers are provided for the captain and first officer.
with the cabin attendant is by an interphone system. A sound hand
set is connected to an outlet in the wheel wells and external servicing
receiver audios are muted during interphone speaking periods with
an interlock to prevent muting of either or both communication audios
during communication periods.
Provision for Additional Facilities
provision is made for the following additional radio navigational
(1) Two VHF navigational receivers
providing omni-directional range and localizer, both installations
having separate controls and instruments for simultaneous
operation. (Magnetic headings for each of the ODR sets derived
from separate remote-indicating compass systems).
(2) Two glide-path receivers
with channel selection automatically tied in with corresponding
(3) An additional marker-beacon
receiver to complete the duplication of radio navigational
of the visual output of either of the two ILS combinations can be
available to the captain by means of one switch. The autopilot automatic
approach equipment would be paralleled with the captains ILS indicator.
has obviously not been possible in this paper to give more than
a bare outline of the work that is necessary in the design of a
enormous amount of test work had to be carried out on the structure,
and functioning of equipment, even before the aircraft first flaw,
and rigorous flight testing is now being carried out to assess control,
stability and general performance.
has been much discussion in the past an the relative merits of jet
and reciprocating engined aircraft, and most of the criticisms of
the jet have been made by people who have never had the experience
of either working on a jet project or really getting down to the
job of comparing the two types on a rational basis.
stage has, however, passed and the main argument now is not if the
jet transport will be used, but when will it be used.
successful demonstration of the C-102 Jetliner in flight has brought
that date a little nearer.
I wish to express my thanks to A.V.
Roe Canada Limited for their permission to give this paper, and
also to the Rolls-Royce Company for permission to publish the engine
My appreciation is also due to Mr.
R.M.Stuart, my Technical Assistant, for his assistance in the preparation
of the art work and diagrams.
A certain amount of the above material
was used in a paper I gave at the Annual Convention of the Engineering
Institute of Canada at Quebec City in May 1949.
Ratio .................................................... 8.31
Ratio at Root................................................ 16.5%
Ratio at Tip.................................................. 12%
of Datum Plans ..................................... 2
on Datum Plane......................................... . 6
Length Overall........................................... 82'
Diameter........ ......................................... 10'
FIG. 1 - Frontispiece.
FIG. 2 - Gust Load Factor Vs, Equivalent Air Velocity for Various
FIG. 3 - Specific Fuel Consumption at Various Speeds.
FIG. 4 - Typical Flight Plan for 500 Miles.
FIG. 5 - Effect of Dive Flaps on Rate of Descent and Comparison
of Rates with Power Off and Half Engine Speed.
FIG. 6 - C-102 Payload Vs. Range.
FIG. 7 - Forty and Fifty Passenger Versions.
FIG. 8 - Nacelle Data.
FIG. 9 - Engine Prior to Installation.
FIG. 10 - Engine Installation.
FIG. 11 - Engine Showing Oil Tank and Gearbox Drive.
FIG. 12 - Variation of Static Thrust with Water-Methanol Injection
at 14,700 R. P. M.
FIG. 13 - Four Engine Take-Off Distance Vs. Temperature at Various
FIG. 14 - Three-Quarter Front View of Nacelle.
FIG. 15 - Three-Quarter Top View Showing Aft Lines of Nacelle.
FIG. 16 - Fuel Consumed During Descent with All Engines at Half
FIG. 17 - Fuel Tank Access Panel in Front Spar.
FIG. 18 - Empennage Showing Double Surface Controls.
FIG. 19 - Cabin Insulation Installation.
FIG. 20 - Flight Deck.
FIG. 21 - Main Undercarriage.
FIG. 22 - Nose Undercarriage.
FIG. 23 - Accessory Compartment.
FIG. 24 - Radio Racks.