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Jim Floyd:RAeS Lecture

Jim Floyd:
RAeS Lecture pg 2

The Arrow Truth!

This republication has been made possible thanks to the assistance of
The Royal Aeronautical Society and Dr. James C. Floyd. This is quite a lengthy lecture and was presented in December 1958. At that time the Arrow was in phase one flight tests.
We hope you enjoy this piece of aviation history.
Scott McArthur. Webmaster, Arrow Recovery Canada.


The Fourteenth British Commonwealth Lecture

The Canadian Approach to All-Weather
Interceptor Development

by

J. C. FLOYD, A.M.C.T., P.Eng., F.C.A.l., M.I.A.S., F.R.Ac.S.
(Vice-President, Engineering, Avro Aircraft Limited, Canada)

The Fourteenth British Commonwealth Lecture," The Canadian Approach to All-Weather Interceptor Development," by Mr.J. C. FLOYD, A.M.C.T., P.Eng., F.C.A.l., M.I.A.S., F.R.Ac.S. was given in the 9th October 1958 at the Royal Institution, Albemarle Street, London, W.1.
The Chair was taken by Dr. E. S. Moult, C.B.E., Ph.D., B.Sc., F.R.Ae.S., Vice-president of the Society, deputising for the President, Sir Arnold Hall, M.A., F.R.S., F.R.Ae.S., who was ill.
Dr. Moult first read a telegram from the President and then introduced the Lecturer, a distinguished Canadian engineer, for this Fourteenth Commonwealth Lecture. Mr. Floyd joined A. V. Roe and Co. Ltd., at Manchester, as an apprentice in 1929, progressing through the design and production offices to become Chief Projects Engineer in 1944. Immediately after the War he joined A. V. Roe Canada Ltd., at first as Chief Technical Officer, becoming Chief Design Engineer in 1949, Works Manager 1951, and Chief Engineer in 1952. He is now Vice-President, Engineering, Avro Aircraft Ltd. Mr. Floyd became a naturalized Canadian in 1950 and in the same year was the first non-American to receive the Wright Brothers Medal, which was awarded for his contributions to aeronautics, including his design of the Avro Jetliner. More recently, he had been known for his work on the Avro CF-100 interceptor and for the Avro Arrow, which made its first flight in March 1958.

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Go Up

(Fig. 2)

NORMAL COMBAT GROSS WEIGHT 64,000 LBS.
WING AREA 1,225 SQ FT. LENGTH 77'9.65"
SPAN 50' HEIGHT 21'.3"

 

THE CONFIGURATION

  There are a number of relatively unconventional features on the Arrow and a reasonably detailed appraisal of these might easily fill a volume of 500 to 600 pages. Therefore, I intend to pick out a few of the highlights and present a broad-brush picture of the design philosophy behind them.
  The RCAF had established a requirement for a two-place, twin-engined aircraft. Their preference for a crew of two was partly based on the complexity of the newer fire control systems and the fact that, while the chosen system was intended to be entirely automatic during the midcourse and terminal phases of the attack, it was the intention to press home an attack on the
basis of a manual mode, in the event of the failure of the automatic mode.
  The choice of two engines was based on a combination of circumstances, the advantages of two engines being obvious in reduced attrition, especially during training. One of the most important factors, however, was the fact that with the very large weapon package required as payload, and the large amount of fuel carried for the range requirements, the size of the aircraft was obviously going to be such that there was no single engine large enough to power it.
  The configuration of the basic fuselage was determined almost entirely by the two-seat, two-engine arrangement and the large armament bay. I will deal more specifically with these items later.

CHOICE OF WING DESIGN

  At the time we laid down the design of the CF-105, there was a somewhat emotional controversy going on in the United States on the relative merits of the delta plan form versus the straight wing for supersonic aircraft.
  We tried very carefully not to become inhibited by association with either side and our choice of a tailless delta was based mainly on the compromise of attempting to achieve structural and acroelastic efficiency, with a very thin wing, and yet at the same time, achieving the large internal fuel capacity required for the specified range.
  This established our delta plan form and the lack of a tail can be attributed almost entirely to our desire not to have to face the problem of putting a tail on top of an extremely thin fin out of the effects of wing downwash, or, otherwise, having to put it so low, again out of the downwash region, that our landing angles would be impossible. We felt that the problems associated with a tailless delta were more predictable and manageable.
  We were also very conscious of the problems that tailed deltas were having at that time, where a large increase in downwash at the stall made the tail strongly destabilizing, so that the stalling characteristics became objectionable.
  It was obvious from the outset that to give the RCAF an aircraft with flexibility of development, the aerodynamic characteristics should be such that they would not limit the speed to less than the structural limitations. The aluminum alloy structure which we favoured was good for speeds greater than a Mach number of 2, and we therefore felt that our aerodyna- mics should be at least as good as this.
  To achieve this we had to go to the thinnest possible t/c ratio, and started out with a 3 per cent t/c wing throughout the span, but aileron reversal forced us to go to a thicker and stiffer wing section, and we compromised at 3.5 per cent at the wing root, and 3.8 per cent at the tip. The structural advantages of the delta made the achievement of a thin wing section possible without a large weight penalty.
  So for us, the tailless delta had distinct virtues, with the added advantage that Avro Manchester had, by that time, done considerable flying with the 707 delta research aircraft, prior to the design of the Vulcan tailless delta bomber, and this information was, of course, available to us.
  However, the delta, like everything else, also had its vices. For instance, aeroelasties were obviously to play a very big part in our design, due to the extremely thin wing and fin sections and, in calculating the aeroelastic and flutter characteristics of a delta wing the standard semi-empirical methods of analysis would have produced a prohibitively heavy structure if we had used them indiscriminately. We had to examine all types of aeroelastic and flutter problems from first principles, and we repeated these as the design progressed. The establishment of the structural matrix was a very laborious process, most of which had to be done on our digital computers.
  Due to the short elevator arm we were in trouble with trim drag. The high elevator angles required to trim at high altitude increase the elevator drag considerably. We investigated means of reducing this and the most promising appeared to be the introduction of negative wing camber. Camber has the effect of building in some elevator angle without the excessive
control surface drag. The amount of camber chosen, which was 3/4 per cent negative, was that which would give a good compromise between the positive angles to trim at low altitude, and the negative angles required at high altitude.

LEADING EDGE NOTCH AND EXTENSION

  Early in the design stages modifications were made to the original clean wing. These were the addition of leading edge droop,
and the semi-span notch with outer wing chord extension.
These modifications were made as a result of wind tunnel tests at Comell Laboratories in Buffalo on a 3 per cent complete model, sting mounted. The approximate Reynolds number used during the tests was between one and two million. These tests showed that we were getting a pitch-up or non-linearity in the Cm-a curve at moderate angles of attack. This phenomenon is not peculiar to deltas, being common to all swept wing aircraft. In flight it could cause a tightening in the turn.
  Crudely, the condition appears to be caused by vortices which start at the tip and move to the apex of the swept wing. Low pressure air is collected from the fuselage and causes a breakaway outboard of the area covered by the vortex, which is mainly at the trailing edge (Figs. 3&4). This causes the effective
FIGURE 3. Vortex pattx plain wing.FIGURE 4. Vortex p of wing with notch and ext leading edge.
aerodynamic centre to shift forward, giving a "pitch-up" or an abrupt change in moment curve.
  While the pitch-up appeared on test to be of small magnitude, since very moderate amounts of pitch-up can be embarrassing to the pilot, attempts were made eliminate it.
  We were aware of the work that had been done by N.A.C.A. and the R.A.E. and the fact that a number of other aircraft which had exhibited this tendency used either notches in the leading edge at about semi span, or extensions of the wing leading edge outboard, in an attempt to prevent the flow separation. The notch had been used, for instance, on the English Electric F-23, and the leading edge extension had been installed on a Grumman F9F9, and a Chance Vought aircraft. The notch has a somewhat similar effect to a fence causes the disturbing vortices to move away from apex of the swept wing toward the notch, which is at semi-span, and reduces the area of disturbed flow over the wing. The notch, however, produces these effects by air flow rather than as a physical barrier. It was our opinion that the effects of the notch are present over the whole speed range, whereas a fence is usually only effective over smaller speed ranges, and the notch was expected to increase the drag by a lesser amount than a fence.
  We did find, however, in our tests that with the notch alone the test results were not repeatable; in other words, the same results could not be got in subsequent tests. When the leading edge extension was installed in addition to the notch the results were far more repeatable. Eight different notches and three extended leading edges were tried in various combinations. The depth of the notch appeared to be the most critical parameter, and here again we had to bear in mind that we could not have too deep a notch because of structural problems.
  Figure 5 shows the effect of the 5 per cent notch, 10 per cent extension of the local chord on the outer wing, which was finally adopted, against the unmodified 31/2 per cent wing at M=0.9, and at an elevator angle of - 20deg.

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