ADA:Avro Jetliner C102-North America's First
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AVRO
C102 Jetliner
North America's First
1949-1956
This
republication has been made possible thanks to the assistance
of
The Society of Automotive Engineers and Dr. James C. Floyd. This
is quite a lengthy lecture and was presented in January 1950. We hope
you enjoy this piece of aviation history.
Scott McArthur. Webmaster, Arrow Recovery
Canada
CONCEPTION
                During
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.
                It
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.
                In
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.
                 Preliminary
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.
 DESIGN
POLICY
                 In
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 conventional
lines.
                 lt,
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.
                 As the
less conventional features will obviously be of the
most interest, these will be covered in greater detail
in this paper.
 SPECIFICATION
The
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 crew.
(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.
(5)
Controllability 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 aircraft.
(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 version)
2,000
miles
Still
air range (Short range version)
1,400
miles
Cruising
speed at 30,000 ft. and 60,000 lb. gross weight
450
+ mph
Payload
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
10,500
lb.
Payload
for 500 mile range with full A. T. A. allowances
at 60,000 lb. T. 0. gross weight
12,000
lb.
4
Engine take-off over 50 ft. obstacle at 60,000
lb.
I. C. A. N. conditions sea level
3,100
ft.
3
engine take-off with above conditions
3,525
ft.
Distance
to Accelerate to Critical Engine Failure Speed
and Stop-ft.,
 
(C.
A. R. 04B.1221):
 
60,000
lb. Gross, Weight at sea level
 
I.
C. A. N. conditions
3,750
'Hot
day'
4,200
       
Landing
Distance
from Height of 50 ft.- ft.
Sea
level (I. C. A. N.)
2,867
3,500
ft. (I. C. A. N.)
3,064
Stalling
speed at landing weight of 50,000 lb.
with flaps in landing position.
87
mph
Stalling
speed at landino, weight of 40,000 lb.
with flaps in landing position.
78
mph
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.
 PRESSURIZING
REQUIREMENTS         
                         To
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.
                           In
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.
                  Since
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.
                   Most
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 pressure.
                   Assuming
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.
                  It
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.
                  As
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 maintenance.
                 Rapid
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.
             As
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
           Originally
designed as a twin engined transport, the C-102 was
designed to take two Rolls-Royce Avon engines.
             In
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.
                It
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. 
              As
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 stage.
                The
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.
                 The
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.
STRUCTURAL
REQUIREMENTS
                 The
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,
                 The
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 gust factor.
                 The
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.
                 The
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.
                To
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.
                The
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 stiffness.
                Extra
heavy skins are used on the lower portion of the fin
to give torsional rigidity and prevent tail flutter.
                 The
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.
                 The
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
Drag
             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.
                 As
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.
                  In
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.
Wing
Section
                  The
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.
                  It
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 landing performance.
                  The
structural problems with high gust factors and the large
amount of fuel which had to be carried also influenced
the wing design.
               The
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.
                This
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
                   A
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 portion.
             It
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.
            Split-type
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.
            The
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.
             The
dihedral on the outer plane is 6 deg., and the wing incidence
is 2 1/2 deg., throughout the span.
Fuselage
Shape
              The
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.
          Special
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.
           A
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.
         The
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
           The
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
            The
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.
          The
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
           The
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.
            The
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.
FLIGHT
PLAN
             Until
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.
            It
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.
                This
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.
                Due
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.
                If
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.
                Any
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.
                 The
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.
                 It
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 377 miles.
                 The
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.
                 Descent
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.
                 As
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 services.
                 There
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.
                 As
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.
                  The
average fuselage angle during descent is not more than
8 deg, which is considered to be reasonable from
a passenger comfort standpoint.
"Copyright
1951 Society of Automotive Engineers, Inc. This paper
is published on this web-site with permission from the
Society of Automotive Engineers, Inc. As a user of this
web-site, you are permitted to view this paper on-line,
download the pdf file and to print a copy at no cost
for your use only. Downloaded pdf files and printouts
of the SAE paper contained on this web-site may not be
copied or distributed to others or for the use of others."
CONVERTED
TO HTML, AND HYPERLINKS ADDED, January 17, 2002.
Scott McArthur.
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