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
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.
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.
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
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.
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
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.
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.
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.
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."
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