Journal of Aeronautical History Paper No. 2015/03

On the Early History of Spinning and Spin Research in the UK Part 2 : The period 1930 - 1940

B J Brinkworth Waterlooville, Hants, UK

Abstract

In Part 1 (Journal of Aeronautical History, Paper 2014/03) a continuous thread of experience and research since WW1 was traced, which had provided a growing understanding of the propensity of an aircraft to spin and a standard procedure that gave the pilot a good chance of being able to recover if one developed. The 1930s were a period of rapid change in air defence policy and technology, leading to the emergence of the monoplane as the preferred configurat- ion for fighters. During this time there was an expansion of work on spinning, with particular reference to progression to the fast flat spin, from which recovery was usually impossible.

A major step forward came with the opening of the RAE vertical Free Spinning Tunnel. Systematic investigations of spinning behaviour could now be made with models that were correctly scaled dynamically. From the results of these and the growing body of tests at full scale, a procedure was devised that distinguished between physical characteristics of aircraft that could be recovered from a spin and of those that could not. Although empirical, this was firmly founded in theory, as the work had been from the beginning.

This procedure allowed designers to include for the first time routine checks at all stages of the design to estimate the likelihood that a new type could be recovered from a spin. The procedure would be refined and become standardised in later work, but the stage reached around the beginning of WW2 marks a distinctive end-point for the early part of the history of spinning in the UK.

1. Introduction

The scientific study of spinning had started in 1917, with a theoretical account of the dynamics of an aircraft in a steady spin, together with the first reported observations of the motion in flight (see Part 1, paper 2014/03). Reviewed at the start of the 1930s, it could be said that work during the intervening years had materially consolidated that position. The critical influence of the moments of the aerodynamic forces acting on the aircraft had soon been recognised. These moments could not yet be calculated, but measurements with models, latterly using the ‘rotating balance’ in the wind tunnel, had begun to give more relevant values of their magnitudes and a general understanding of their effects. In particular, it was revealed that any displacement in yaw of the aircraft relative to its helical path was of first importance in determining the required moments, though that had been neglected in earlier work.

Flight-testing at full scale would always be required, to expose new problems and to validate conclusions reached from theory and testing with models. A standard instrument package had been developed for this, which could now record automatically all the measurements

168

Journal of Aeronautical History Paper No. 2015/03 required to define the behaviour of an aircraft in the spin. It had also been understood that the outcome of research needed to be communicated more readily to personnel in the industry, where designers were becoming increasingly receptive and better placed to implement advice when it became available.

Yet in its Technical Report for 1929-30, the ARC was obliged to state that ‘Existing knowledge is insufficient to prevent the occasional appearance of an aeroplane upon which it is difficult, or impossible, to check a fully developed spin’ (1). And in the Supplement to that report, it had gone on to endorse the opinion of its Air Ministry member that ‘the spinning of aircraft presents one of the most important practical problems yet to be solved’. Accordingly, the Committee had recommended that research on the subject should ‘take first priority’.

From a purely practical viewpoint, it was necessary that much of the work on spinning would continue to be done at RAE and NPL with models. It had been realised that the aerodynamic environment of an aircraft in a spin was unique, and for measurements on models to be useful the situations in which they had been obtained had to represent that environment closely. This would be a leading theme in the development of experimental methods in the following years. There were continuing concerns about recovery from spinning, particularly from the flat spin, characterised by very high angles of incidence and high rates of rotation. Unexpectedly large moments were being required from the controls in both the pitch and yaw directions to effect recovery.

The spinning trials of service aircraft at A&AEE Martlesham Heath and MAEE Felixstowe would continue. In this Annual Report there was a specific mention of further tests to be made on the Fairey IIIF aircraft, for which the spin had ‘sometimes been a very unsteady motion’. This single-engined aircraft was a versatile machine that served in a variety of roles for both the RN and RAF over a service life of 14 years. Unsteady motion usually consisted of nutation (a nodding movement) imposed on the rotation, and there was a belief that this indicated that the aircraft was close to moving on to a flat spin. The Fairey IIIF could be fitted interchangeably with wheels or floats. There was concern about the possible effects of the proportionately large contributions made by floats to the moments of inertia and the base and side areas, which are apparent in the photograph of Figure 1. The case shows the complexities likely to be faced in reaching reliable judgments about the spinning characteristics of a given aircraft type at the time. It had not yet been possible to do that fully before the aircraft had been required to spin during trials for acceptance into service, and modifications made to it afterwards could also be problematic. Figure 1 The Fairey IIIF multi-role aircraft, with floats fitted (National Aerospace Library Collection)

169

Journal of Aeronautical History Paper No. 2015/03

The wings of the Fairey IIIF had been set at zero stagger, a feature known to be implicated in progression to the flat spin, but the gap between the two wings was large enough to be a signif- icant mitigating factor in countering that. To establish a numerical base for an investigation, initial measurements were made of the rolling and yawing moments due to roll with models at 1/15th scale, both with and without floats, in the rotating balance at NPL (2).

The rolling moment about the spin axis due to the fin and rudder was negative (against the spin) at all incidences tested, except for the highest value of 61o, a value that could be reached in a flat spin, when the moment became positive. However, the moment due to the addition of floats was positive at all incidences. The yawing moment about the body axis due to the floats was also negative over most of the incidence range, but it too became positive at the highest incidence tested. Measurements without rotation were made of the forces and couples for all body axes, reflecting the considerable variations with incidence and sideslip that had been noted regularly in other cases previously.

The opportunity was taken in these tests to build up further basic information on the effects of changes to the tailplane shape and position and effects of using differential and floating ailerons and ‘interceptors’ (shallow strips that could be raised from the upper surface of wings, usually in association with tip slots, although the slots themselves were not represented on the models tested on this occasion).

When used in calculations, these results gave quite good agreement between the calculated values of incidence and spin rate and the full-scale spin results already available for the landplane version of the aircraft. Though the margin of safety was lower for the seaplane version, it was indicated that there should be no great difficulty in recovering it from the spin. However, during the trials made at MAEE at Felixstowe the testers had been unable to do so at one point and had to abandon the aircraft, so in service the crews were advised to do that at once if a spin developed or continued below 1,500ft. This would be a region in which seaplanes would commonly be operating, but it appears that no further work was done on spinning with that version of the type.

More generally, progress in understanding spinning was a result of interaction between theoretical analysis, experiments in wind tunnels and flight tests. As such, it is an excellent example of the value of this combined approach to aeronautical problems.

2 Communication with designers

During the 1920s, UK industry had not followed the US in raising its productivity sufficiently to be fully competitive in world markets. This situation was to continue into the 1930s due to national monetary policies which over-valued the currency, while at home it was difficult to obtain the finance necessary to modernise plant and equipment. Stringent economy was necessary for aircraft firms to remain in business, but some changes did occur that contributed to an increasing awareness of the importance of what was being found in the various aspects of aeronautical research and to apply that to design.

170

Journal of Aeronautical History Paper No. 2015/03

At this time the number of persons in the industry with full professional qualifications was still quite low. The design team for a new project would continue to be a small group, affording few opportunities for ambitious newcomers to enter. Although a modest proportion of senior people were members of the RAeS, many had risen under the pressures of WW 1, largely on a basis of the hard lessons of experience. There was only a gradual realisation of the need for higher competence in engineering science as a basis for commercial success in this field. Where for example a recent graduate was taken on, it would at best be in a lowly role as an assistant to one of the principals in the design team. Opportunities to show initiative were few and advancement was slow, but later memoirs would show that some of those who were able to contain their frustrations at that time were to become the more significant figures in the industry as they gained seniority.

A lecture on spinning to the Yeovil branch of the Society by S Scott-Hall, then of A&AEE, was typical of the continuing support given to professionals, despite the contracted state of the Air Ministry establishments, and provides a convenient illustration of the situation as it stood at the start of the new decade. The text of this lecture was published in the Journal of the Royal Aeronautical Society for 1931, though oddly without its illustrations (3). Those attending, and later readers of the paper, were first warned that the spin must still be taken very seriously. Whatever research had been done and would be done in the future, any spin must be considered to be hazardous. The acceleration experienced by the pilot in the rotation would not seem large enough to be troublesome, but even the most experienced could quickly become disorientated by the motion. If he was in cloud, the pilot might not even realise that a spin was happening, but in clear air with plenty of visual cues, it had been found that the pilot could become so confused as to be unable to tell even in which direction he was rotating. If the spin was prolonged he was likely to become nauseated also. This was made worse when the motion was unsteady, where substantial periodic oscillations in nutation could be superimposed on what was otherwise a steady rotation.

Firms engaged only in the production of civil aircraft should not think of spinning as irrelevant to them. When there was any possibility that aerobatics would be undertaken, the official responsibility for issuing a Certificate of Airworthiness for the aircraft lay with A&AEE, where nearly the same investigation of proneness to spin would be included as on aircraft being evaluated for possible service with the RAF. All aircraft were first required to be stabilised for eight turns after entering the spin. Then to satisfy the criteria for acceptance, military types must come out in not more than three further turns after moving the controls to begin recovery, and civil aircraft in not more than four turns. Spin tests were made in both directions, as the motion was affected by the angular momentum of the engine and airscrew, even when at idle. So far, test protocols had been established only for single-engined aircraft, though it had been conjectured that spinning could be more troublesome in some twin-engined types. This would perhaps result from the different distribution of mass in these, tending to change the relative moments of inertia about the three principal axes.

Scott-Hall went on to review recent research activity along lines as reported towards the end of Part 1 of this paper. Though theory had revealed the forces and moments that were most strongly involved in the spin, and measurements with the rotating balance had enabled these to be evaluated to some extent, it was still only when an aircraft was spinning at full scale

171

Journal of Aeronautical History Paper No. 2015/03 that it was fully free to respond to all of these acting together. The standard set of instruments developed by RAE for measuring the response in flight now allowed the relevant data to be recorded automatically. Work at full scale progressed slowly and was expensive, but would need to continue.

A good part of the lecture was devoted to features of aircraft that were already shown to increase proneness to spinning and slowness or impossibility of recovery, with detail that would be expected to be of especial interest to designers. For this audience, Scott-Hall showed what could be learnt from interpretation of the three Euler equations for the moments in roll, pitch and yaw. The importance of the moments of inertia was emphasised, with illustrations showing how calculated values of these properties had been checked by measure- ment, suspending an aircraft in various ways and measuring the periods of small oscillations.

Factors affecting the aerodynamic moments available from the empennage were outlined. Designers would now be accustomed to ensuring that moments available from tailplane and fin / rudder, represented by the volume coefficients, were adequate for longitudinal and lateral stability in normal flight. But in a deep stall with sideslip the effectiveness of the empennage was greatly reduced by interference from the wakes from the tailplane and rear fuselage, and in some orientations, the wings also. The ARC had recommended limiting values for the fin and rudder volume coefficients, currently not to be less than 0.08. Practices adopted for reducing the shielding of the fin and rudder by the tailplane and rear fuselage in the spin were reviewed, including some reference to approaches used in the USA.

This comprehensive presentation was followed a year later by another article along similar lines by Scott-Hall in Aircraft Engineering (4), which would also reach the wider audience of technicians and draughtsmen mentioned earlier. Some of the highlights in these presentations will be reported more fully later.

The Society of British Aircraft Constructors Ltd (SBAC; today the Society of British Aerospace Companies) had been formed in 1916, and from the beginning one of its aims had been to increase awareness of new developments across all parts of the industry. In 1932 it approached the ARC to enquire whether it was now possible to formulate ‘a set of simple rules for the use of aeroplane designers to guide them in making their machines safe from spinning’. The composition of such a review was remitted to H B Irving of RAE, who was the principal investigator of spinning there. It would be published as R&M 1535 in March 1933 (18), which is reviewed in Section 5 below.

3 New departures 3.1 Approaches in the early 1930s

There would always be uncertainties about the validity of applying results from wind tunnel tests to the evaluation of aircraft performance, particularly arising from the general influence of what was known as ‘scale effect’. The work of Prandtl and others had shown that the airflow in the boundary layer next to the surfaces of the model could not be fully represent- ative of that on the aircraft. The flow was determined by the Reynolds number, which as well

172

Journal of Aeronautical History Paper No. 2015/03 as factors such as size and speed includes the effects of the viscosity of the air. On the model this number would be typically orders of magnitude smaller than at full scale. As yet, there were few indications about the importance of this. However, in the spin the lifting surfaces and associated controls were operating in deeply separated and highly turbulent flows, where it was thought possible that Reynolds number effects might be less significant. Spinning trials at full scale would continue to be essential, where key quantities could now be measured that allowed more reliable estimations to be made of the aerodynamic moments experienced. This was vital, for providing data by which both the success of theory and the validity of model experiments could be judged, though inevitably it would be a slow process. Meanwhile, refinements of the theory were being pursued, and thoughts were turning towards new experi- mental methods with models, in which all the leading factors involved in spin behaviour could act together, with a greater prospect that the intricacies of that state could be reproduced more closely.

Hitherto, all theoretical work had been on modelling the steady spin, and this aspect was now thought to be fairly well represented. The unsteady phases of entry into and recovery from the spin could also be monitored with automatic recording in flight, but theoretical modelling of these would involve velocity-dependent derivatives of which nothing was yet known. Developments in this area were not to be expected shortly. But at this time, staff at NPL were engaged in a long-term project to investigate the stability of spins, drawing on contemporary developments in other areas of stability and control, by which they hoped to be able to identify and estimate the key derivatives required. An early report from this, concerning the relative effectiveness of the three controls in starting a manoeuvre that would lead to recovery, would appear in due course.

In the area of model testing, significant developments of the rotating balance were being put in hand at NPL that would allow moments to be measured with whole aircraft models in rotation about an offset axis, with the objective of obtaining a closer representation of the full extent of coupling between the actions around all three body axes. Meanwhile, attempts were made in a new direction under A V Stephens at RAE, where the behaviour of spinning models was observed for the first time in free descent. These new approaches were to be exercised immediately in the continuing study of a further case of serious difficulty in recovery from the spin.

3.2 Aircraft ‘H’ and the Hawker Hornbill

Spinning tests with models in the rotating balance and at full scale had been made towards the end of the 1920s concerning a single-seater fighter aircraft identified simply by the designation ‘H’, as reported in Part 1. This is recognisable as the Mk 1 version of the Gloster Gamecock, of which 90 had entered service with the RAF. It had proved to be prone to accidents, in some of which spinning had been implicated. The rotating balance testing had indicated that spinning could be alleviated by lengthening the rear fuselage of the model and fitting an enlarged fin with a horn-balanced rudder. Full-scale tests at RAE on a modified aircraft in which these changes had been approximated showed that the spinning behaviour was significantly improved.

173

Journal of Aeronautical History Paper No. 2015/03

In 1931 a revised form of the aircraft was submitted for testing, probably with a view to preparing a new Mark II of the type that would be acceptable for entry into service. As suggested from the previous tests, an enlarged fin and rudder had been incorporated, though without significant changes to the rear fuselage. In renewed spinning tests at full scale at A&AEE it was found that after a few slow steep turns, the new machine had a tendency to flick rapidly into a fast flat spin, even when the centre of gravity was moved to a forward position (5). Recovery could be effected from spins to the left, but in a spin to the right this proved very difficult. With great coolness, the pilot, Fl Lt C E Maitland, had continued to try various measures to start recovery during an estimated 30 turns at a rate of almost one o revolution per second at an incidence of 57 . Having descended 8,000ft while spinning, he had decided to abandon the aircraft, but his movements in standing up to do so somehow started a recovery and he had managed to regain control and land safely (1). Subsequently, he reported that he could not be certain that his actions had been entirely rational throughout the descent, as he was ‘only in full possession of his faculties towards the end’, when he had ‘got used to the conditions’. Records of rudder position suggested that at times he had been uncertain about the direction required for recovery. Despite this experience, he went on to make a further spin, with the centre of gravity further forward, though again the fast flat spin had developed and he found that applying full rudder against the spin needed a force that was hard to maintain for more than a few seconds. Rocking the controls and opening and closing the throttle had no effect, but when the engine stopped the aircraft finally recovered. This was presumably due to the absence of the additional angular momentum of the engine and airscrew about the longitudinal axis when these were rotating, perhaps combined with the disappearance of the slipstream effect on the fin, which together had brought the recovery within reach. But estimates made by calculation shortly afterwards indicated how marginal the situation had been - the pedal force required to balance the centrifugal effects alone on the rudder at this rate of spin would have been about 100lb (6).

Following these experiences, further tests were made involving different modifications to the original form of Aircraft ‘H’ (7). Changes reported previously had included lengthening the rear fuselage and substantially enlarging the fin and rudder, as shown in Figure 2a. For new model tests in the NPL rotating balance, the original length was retained (as in the revised aircraft), but the after part of the fuselage was deepened and the tailplane raised from near the centre-line to the top of the body. Rolling and yawing moments due to rolling and to sideslip were measured over a range of incidence, on the whole (1/10th scale) model and on the wings and body separately. At the higher incidences, pitching moments and rudder control while rolling Figure 2a Aircraft with very different spinning were also measured. Although characteristics - RAE Aircraft 'H' valuable comparative data were (National Aerospace Library Collection)

174

Journal of Aeronautical History Paper No. 2015/03 gathered on the moments with a deepened and with a lengthened rear fuselage, it was o concluded that at an incidence of around 50 , at which the full-size aircraft had spun, the differences were not great.

There appears to have been wide discussions of the situation at RAE before resuming full- scale testing of Aircraft ‘H’, as the modifications then adopted are said to have been suggested by William Farren, formerly head of Aerodynamics Department, but then at Cambridge (7). The rear fuselage was retained in length and profile, but was deepened in a practical fashion by extensions to the fin forwards along both the dorsal and ventral surfaces. The large rudder with horn balance, a distinctive feature of this aircraft, was also retained, but the tailplane, though unchanged in shape and area, was raised nearly to the top of the fin. The usual instrumentation fitted for the tests provided records of the rudder position and the normal acceleration at the centre of gravity. As shown earlier, the incidence in the spin could be estimated from the recorded normal acceleration alone. However, there was now the novel addition of a pinhole ciné-camera mounted above the top wing on the starboard side, enabling the rate of rotation to be obtained from the film record. If the sun came within the field of view, a further estimate of the incidence could be determined also, by reference to the sun’s elevation given in astronomical tables. This provided a check on the incidence value obtained from the accelerometer reading, together with a way of estimating it if the engine was running, when the usual assumption that the acceleration was normal to the wing chord was no longer valid. The rate of descent was still obtained by the pilot from the altimeter and a stopwatch.

It was found that this raising of the tailplane and increasing the fuselage lateral area had entirely eliminated the vicious spinning characteristics of Aeroplane ‘H’. There had been an appreciable reduction of incidence and rate of turn in the spin, with no creep towards a flat spin, even when the descent was prolonged to 25 or 30 turns. Moreover, rapid recovery could be made from all spins. The opportunity was taken to extend the trials to include studies of the effects of moving the centre of gravity position, varying the engine power during the spin, and use of the elevators at different times during the descent and recovery. These indicated that with the engine running during the spin or at the moment of reversing the controls the machine could be brought out with a smaller loss of height and at a dive angle that was noticeably less steep.

When the tests were concluded with a programme of aerobatics to see if the excellent manoeuvrability of the original machine had been impaired by the modifications made to it, the pilots were of the opinion that its qualities had not just been maintained but enhanced, both in aerobatics and in normal flight. However, the Mk II Gamecock that incorporated many of these changes was not ordered for the RAF, though three were delivered to the Finnish air force and a further 15 were built in Finland under licence, remaining in service there for nearly 20 years.

Not all such investigations were made because of dangerous behaviour, and not all could yet produce such conclusive results. An example of the frustrations that continued to arise in spinning work is provided by the case of the Hawker Hornbill, shown in Figure 2b. When evaluated at A&AEE, it was considered to be ‘an advanced concept in high-speed design’, perhaps because of its sleek profile due to the use of an in-line engine. However, it had been

175

Journal of Aeronautical History Paper No. 2015/03 found unsuitable for service use on the grounds of being under-powered, so that its operational ceiling was below the requirements of the test programme. Its directional stability and control in normal flying conditions were also unsatisfactory. But pilots who had tested the prototype had given very favourable reports about its stability near the stall and its spinning behaviour, where no difficulty had been experienced in recovering from spins with it in Figure 2b Aircraft with very different spinning either direction. It was decided to characteristics - Hawker Hornbill remit it to RAE for further (National Aerospace Library Collection) investigation.

A substantial programme of work was undertaken in 1931/32, aimed at discovering the reasons for its favourable spinning qualities, beginning with model tests (8). Initially, these were mainly made with a model having wings and body only, which unusually was sent to the USA to obtain its lift and drag characteristics in the NACA variable density tunnel, where possible Reynolds number effects might be revealed. The results showed that there was an appreciable scale effect on the maximum lift coefficient, but little change in the shape of the lift curve, which had a considerable region of negative slope at incidences beyond the stall. This was expected to help in stabilising the post-stall behaviour and, by Glauert’s rule, to limit the range of incidence in which autorotation could occur.

More extensive tests were carried out with a 1/12th scale model in the rotating balance at o RAE, at incidences up to 60 and for the parameter ps/V up to 0.9 (a measure of the angle of the spiral traced by the wing tip; notation is given in the Appendix after the references), with wings and body and wings only. These showed that the presence of the body had a marked effect on the rolling properties of the wings, and the yawing moment in rolling with the original body compared favourably with that of the lengthened body fitted experimentally to Aeroplane ‘H’. Spinning calculations made on the basis of these measurements indicated that flat spins should not be possible with this aircraft and recovery should be easy.

Model tests were followed by further flight tests at RAE. For these, the aircraft was equipped with extensive instrumentation similar to that used for Aircraft ‘H’. The installation of this had taken the centre of gravity to an abnormally rearward location, and a slightly forward stick position was necessary to obtain consistent behaviour in the spin. It was now found that the motion in the spin was uneven, with the incidence fluctuating by up to 7 o on either side of the mean value. The stall-stability tests had also shown variations that had not been reported for the tests at A&AEE, though these effects were now attributed to differences in the atmospheric turbulence levels on the two occasions. But as before, recovery was found to be easy and prompt for spins in both directions.

176

Journal of Aeronautical History Paper No. 2015/03

It seems that these extensive tests had served mainly to confirm the difficulty of proving a negative - why something does not happen. While the model tests and calculations had agreed with the flight tests in that spin recovery should be easy, no particular features of this aircraft were identified that could account for its good qualities. It was concluded that they had been due to ‘a happy but fortuitous combination of parts of the design’, but quite what that had been was still obscure.

Even though they would not always produce the evidence hoped for, the outcome of comparative tests of this kind was considered by the ARC to be ‘highly satisfactory’. Further, they had shown the growing maturity of spinning test methods at both NPL and RAE.

3.3 Free spinning with models

Aircraft ‘H’ featured again in a new technique using models in free descent which was being explored at RAE. Correctly scaled in physical shape, mass, moments of inertia and centre of gravity position, these were released indoors from a walkway in the apex of the roof of the Balloon Shed. A drop of 80 ft in still air was available before the models were caught in a net spread close to the floor (9).

At first, the models were launched in a stalled attitude from a swinging trapeze, with the control surfaces set to produce a spin. For investigating spin recovery, a simple device using an air bleed was fitted, that moved the controls to the required positions after a suitable time delay. However, the height lost in establishing the spin state meant that only a few steady rotations could be obtained before the model reached the net. With the stabilised spin limited to such a short period towards the end of the descent, there was uncertainty about whether it could correctly be assumed to have become steady and gave little time for recovery actions to act, so a different method of launching the model was devised. In this, it was given a suitable combination of initial orientation, vertical velocity and rotation by the apparatus shown in Figure 3. The model was mounted on a Figure 3 Rig for launching a dynamic model in carriage that could slide down and round spinning attitude for free descent (Aircraft ‘H’) a vertical rod with a thread to produce (Reference 9)

177

Journal of Aeronautical History Paper No. 2015/03 an appropriate helical motion. Three rods were available, with spiral pitches giving a range of rotation from 0.26 to 1.04 radians per foot descent. A trip released the model after a fall under gravity of nearly 7ft. Most tests were made with the model in the initial orientation of the prototype spin, but in some it was arranged that there would be no sideslip, by suitably adjusting the position in yaw.

With the start provided, free spinning was induced successfully at a much earlier stage than before, and a systematic series of tests could begin. Records of the attitude of the model were extracted from films taken with a ‘cinema camera’, initially with the model followed by the beam of a 36in searchlight, and later in silhouette against the sky, a background provided by opening the full-height door at the far end of the building. Close-up filming confirmed details of flight experience that, although spins were steady in the sense of being repetitive, there was often significant nutation superimposed on the rotation.

The model shown on the launching apparatus in Figure 3 is that of aircraft ‘H’. Launched in its original form it immediately took up a flat spin of the worrying kind found at full scale. Trials were then made with modifications that located the tailplane progressively higher, up to a position near the top of the fin. The recovery was increasingly improved, and a further enhancement was obtained by deepening the rear fuselage, as suggested by previous work.

Other models used in this first free-descent testing had been of the Bristol Fighter, a type notable for its manoeuvrability and popular with pilots, and now virtually a standard case in spinning trials, which had earlier shown it to be subject to irregularity in the spin. It was soon found that the model would take up one or other of two distinct spin states, as had been observed at full scale. When launched at a moderate incidence, the spin was of the common o slow, steep variety, with a mean incidence of 33 and rotation at 9rad/s, but with higher initial o incidence a shallow, fast spin resulted, with mean incidence of 65 and rotation at 15rad/s. There were occasions when a transition occurred from the slow spin to the fast one during the descent. Thus, in these tests, models were being seen for the first time to replicate behaviour that had been experienced in tests with the full-scale aircraft. The effects of systematic variations in the test conditions could now be interpreted with more confidence.

Models were fitted with variable and moveable weights, allowing the position of the centre of gravity and the moments of inertia to be altered. Moving the centre of gravity further aft, which in practice would reduce stability but increase agility in manoeuvring, caused an increase in incidence in both steep and flat spins, and made the latter more likely to occur. Other tests were made with weights attached to the wing tips, so as to double the moment of inertia about the longitudinal axis (with an addition to that about the normal axis also), causing a marked increase in the incidence in the flat spin, with a faster rotation and slower rate of descent. There was little effect on the mean orientation or rate of rotation in the steep spin, but there were noticeable fluctuations of these quantities about their mean values, which were suspected of being evidence of an increasing likelihood of transition to the flat spin. Changes in the spinning characteristics caused by adding weights at the extremities of the fuselage, to double the moment of inertia about the lateral axis (again with an increase about the normal axis), were generally in the opposite direction to those caused by increasing the inertia about the

178

Journal of Aeronautical History Paper No. 2015/03

longitudinal axis. Systematic variation of test conditions, with results of such utility, had not been possible in this field before.

Testing in free-descent also allowed practical measures for improving recovery from the spin to be investigated promptly with models. Since it was accepted that the first action in effecting recovery should be to apply rudder to stop the rotation, the measure chosen was usually one that would reduce shielding of the fin and rudder at high incidence by the wake from the tailplane and fuselage. Six changes to arrangements of the empennage and rear fuselage were made to the Bristol Fighter model, to compare the spin recovery with that of the standard form. The results again showed clearly that the most effective changes were to deepen the rear fuselage and to raise the tailplane relative to the rudder, including placing it at the top of the fin. It was also confirmed that a tapering rear fuselage of circular cross-section was significantly disadvantageous, as had been found at full scale in the problematic Springbok and Bantam aircraft, reported in Part 1.

The overall outcome of this early experience with free spinning was considered to be very promising. As well as adding greatly to empirical knowledge of the influence of the configuration of aircraft on their spin characteristics, which had not been so accessible with previous methods, effects calculated from theory could also be checked with more confidence. Spin recovery techniques and the effects of the size and positioning of the empennage on recovery were now within reach of systematic investigation with models. However, at this point the free-descent tests had to be terminated through the implementation of a decision to demolish the old Balloon Shed that had been made at the beginning of 1931. Yet it could be said that they had already sufficiently served their purpose. It had by now been concluded that it would be necessary to find a way to prolong the free-spinning process so that there could be full confidence that the state had become steady when measurements of its charact- eristics were made and recovery actions commenced. And so attention turned to the method of using a vertical wind tunnel, in which (with the air speed correctly adjusted to equal the rate of descent) the model could be kept in suspension, spinning in free flight but stationary against its background for as long as might be needed to make the required measurements.

3.4 Vertical wind tunnels

The idea of using a vertical wind tunnel for spin investigations may have occurred around this time to different people. Among the early users in the UK was W E Gray, though as he was at the time a freelance investigator, the exact date of his experiments is unknown (said to have been made initially in a small tunnel working with air flow from a domestic vacuum- cleaner). But a legacy of Gray’s work is available in another vertical tunnel of his, now in the reserve collection of the Science Museum, as shown in Figure 4. This perhaps dates from early in 1932. Of very basic construction, about 7ft high with a working section of only 15in in diameter, it had a maximum air speed of 25ft/s. Air was drawn through the tunnel by a fan at the top, driven by a 75W motor, entering through a flared inlet in the base, and after passing through the working section, was slowed by a diffuser as it approached the fan. Speed adjustment was made manually, via the rheostat visible at the top. Gray is believed to have used it first to investigate a matter of his own interest, the behaviour in the spin of

179

Journal of Aeronautical History Paper No. 2015/03

have used it first to investigate a matter of his own interest, the behaviour in the spin of aircraft

Figure 5 Model of Mignon Pou-du-ciel for Gray's spinning tunnel tests (Science Museum Reserve Collection)

Figure 4 Gray’s 15in diameter vertical spinning tunnel (Science Museum Reserve Collection) aircraft with the ‘V’ (or ‘butterfly’) tail. Only very small models, typically of 4in span, could be tested in this tunnel, and one of these models, shown in the enlargement of Figure 5, has survived and is preserved with it. This was of Henri Mignon’s tiny sports aircraft Pou-du-ciel, and the close-coupled tailplane of its tandem configuration is evident. It had been claimed that with that layout the aircraft could not be stalled, and therefore would not spin. Gray’s results seem not to have been published, though it is said that he was able to get the model spinning readily enough.

The introduction of vertical tunnels at RAE is better reported in the literature. Before building the much larger tunnel that would Figure 6 RAE 2ft diameter spinning tunnel be required in practice, one of 5ft height and for proof-of-concept tests 2ft diameter was constructed in 1931, (FAST Museum collection)

180

Journal of Aeronautical History Paper No. 2015/03 octagonal in cross-section so that the model could be viewed from many directions through plane glass walls. Shown in Figure 6, this was the first vertical tunnel of the time that could be dated, though it was intended primarily to prove the concept, rather than to be a tool for research. An account by Stephens of work leading up to it was published in Aircraft Engineering in September 1931 (10).

As with Gray’s tunnel, an electrically-driven fan was mounted at the top, with a flared entry in the base, followed by a honeycomb straightener to assist in providing a uniform flow at entry to the working section. There was no diffuser at the exit, as used by Gray to reduce the air velocity there and prevent the model from moving up into the fan. A vertical spigot was fitted at the centre of the honeycomb, on which a boss in the model could be loosely placed, presenting it to the flow at a typical incidence for the steep spin. When the fan was started, the model would begin to rotate, and on increasing the speed, it would be lifted off the spigot to take up quickly the characteristic motion of the spin. Tests confirmed that the model remained close to the axis of the tunnel, and could be readily held in a hovering position by small adjustments to the tunnel speed. This was measured to give the equivalent value of the rate of descent in still air. Small models of up to 6in span were used, but proved sufficient to check the basic principles of this mode of testing. As expected, the rate of rotation of small models was fast, but could be measured with a stroboscope. Models of the same aircraft that had been tested in free flight were tested in this tunnel, though with the small scale required, exactly equivalent mass distributions could not be obtained readily. However, the results gave a clear indication that similar modes of spin were observed in the tunnel as in the Balloon Shed tests.

This small tunnel was taken to the Lecture Theatre of the Royal Society of Arts in November 1931 to provide illustrations for a general lecture on recent work on spinning given by Irving and Stephens (11). This was quite a wide-ranging survey, on the theme of safety in spinning, with particular reference to factors that could lead to the flat spin, from which recovery had been difficult and often impossible. The character of a spin was described, with an explanation of the interactions between the inertial properties of the aircraft and the aerodynamic character- istics experienced in the rotating motion of the spin, as reviewed here. It was explained how the effectiveness of the fin, in limiting the tendency to move towards a flat spin, could be greatly diminished as a result of shielding by the wake of the tailplane in the deeply stalled and yawed orientation. The action of the rudder, which was the principal agent for recovery from the spin, was similarly vitiated. Recent theoretical work to clarify these effects and experiments to measure some of the moments and their derivatives with rotating balances were described. Some results of spinning tests at full scale and those so far made with models in free flight and in the vertical tunnel were outlined, giving a clear impression that understanding of the spin was progressing over a wide front. A note was added to the account of this lecture that appeared in the Journal of the Royal Aeronautical Society (11), in acknowledgement of the contribution of Mr K V Wright in recent years, giving him credit particularly for methods of making measurements in spins at full scale and with models and for originating the technique of the free flight model experiments.

Trials with the small tunnel provided adequate justification for the construction of the RAE Free Spinning Tunnel at the western end of Beta Shed, one of its taller buildings (No 29), as

181

Journal of Aeronautical History Paper No. 2015/03 outlined in Figure 7. Opened in November 1931, this had a circular working section 30ft high and 12ft in diameter, with a four-bladed fan driven by a directly-coupled 50HP DC motor, giving a maximum tunnel air speed of 35ft/s. Sensitive speed control was obtained through the use of a Ward- Leonard arrangement, which employed a rotating power amplifier in the form of a coupled motor-generator set, the motor running this at a constant speed from a separate supply. The operator obtained fine control of the fan power by varying the voltage applied to the field windings of the generator, which in turn supplied control power for the fan motor from the armature windings. An airtight observation room was built into the side of the tunnel, from where the tunnel speed was controlled manually and the motion of the models could be observed directly and recorded. Though the opening from the gallery into the tunnel was 4ft high, it was found not to affect the flow distribution significantly.

In this large tunnel, models of around 2 Figure 7 Arrangement of the RAE 12ft Free to 3ft span could be flown comfortably, Spinning Tunnel (RAE Report Aero 2456) as illustrated in Figure 8 (12). This shows how the model was introduced into the tunnel from the gallery of the observation room at the end of a swinging arm carrying a tall spigot, on which it was mounted with a suitable initial attitude as in the small tunnel trials. Once the fan was turned on and the model had risen from the spigot, the arm was swung back into the gallery. The rotational speed could be obtained with sufficient accuracy by counting against a stop- watch, but for the final record this and the angles of bank, pitch and yaw Figure 8 View from the observation gallery of the were obtained from ciné-camera film. RAE 12ft Free Spinning Tunnel (Reference 12)

182

Journal of Aeronautical History Paper No. 2015/03

For experiments on recovery from the spin, the controls of the model were activated at first by a spring-loaded dashpot with a variable air bleed, allowing ample time to ensure that the spin state was steady before recovery was initiated. Later, clockwork mechanisms and radio control were used for this. Models coming out of the spin were generally in a steep dive, and were arrested on a net which extended across the tunnel at a level somewhat below that of the observation room.

The Free Spinning Tunnel proved to be an asset of great utility, requiring little in the way of modification beyond the fitting in 1938 of a more powerful fan motor to obtain the increased air speeds required for modelling of aircraft with increasing wing loading. It remained in use for several decades (on one occasion by the present writer). In its Technical Report for 1931/32, the ARC recorded that it ‘attached much importance to this line of research’ (12).

3.5 Scaling factors in the spin

Once it became possible to measure quantities from a model in a free spin it was necessary to establish the rules by which the observed values could be scaled up to form comparisons with measurements on the full-size aircraft. At first, this was done by comparing inertial forces, 2 giving rise to simple scaling rules derived from assuming that groups such as V /sg would have the same values for the aircraft and its models. However, it was shortly realised that this would not take account of a difference in air density between the model test at ground level and the aircraft spinning at altitude. Initially, some quite complex arguments were used to account for the effects of altitude on the spin, but in due course a straightforward procedure came into use. This is based on finding the ratios of quantities for the aircraft and its model, as developed below from the basic equations of motion for the prototype spin case from Section 4.1 of Part 1 (pages 136 to 139 of paper 2014/03). Notation is given in the Appendix at the end of Part 2 (this paper)*.

Given that in the spin the drag is equal to the weight 2 2

W = D = CD ½ ρ V S or V =2W/CD ρ S (1) Then if quantities for the aircraft take the suffix a and for the model m, the ratio of their rates of descent is given by 2

(Va/Vm) = (W/CD ρ S)a/(W/CD ρ S)m (2) To obtain straightforward scaling laws, expressions such as equation 2 were simplified by making the basic assumptions that: a) if investigating spinning near the ground the model would be made to the same average weight per unit volume as that of the aircraft. This is proportional to the group W/Ss. As the density of the air was the same for the model and the aircraft, quantities on the right hand side of equation 2 would cancel, leaving only the ratio

(CDs)a/(CDs)m ; b) if the test is made at ground level but is intended to simulate a spin at altitude, the effect would be taken into account similarly by making the model to that density

* Terms not defined in the Appendix are W, weight, and s, semi-span

183

Journal of Aeronautical History Paper No. 2015/03

which was in the same ratio to the air density when on test as for the aircraft at its operating height. With no significant difference in gravity with altitude, this required equality of the group W/ ρ Ss, which became known as the ‘weight density ratio’.

With equality of that, the right hand side of equation 2 again becomes (CDs)a/(CDs)m; and thirdly

c) coefficients such as CD would be similar for the model and the aircraft. This was an assumption that could be validated only by experience. In conventional wind-tunnel work, there was evidently a ‘scale effect’ in the results, but in the spin the wings and tailplane are deeply stalled, and the flow 3-dimensional and highly-separated. In the current state of knowledge it was then a fair initial assumption that differences in the boundary-layer flows would not have much effect and the drag coefficients for the aircraft and its model would be similar.

If the three assumptions are valid, it follows that with a linear scale of 1 to n for the model (for example n = sa/sm), equation 2 gives the first scaling law

Va/Vm = √n (3)

To scale the model correctly, the density of the air at the altitude of the aircraft relative to the sea level value must be represented. For example, routine spin testing at A&AEE Martlesham Heath began at 15,000ft at this time, where the relative density σ is about 0.63*. Other values were required to be assumed when the behaviour at different altitudes was being investigated more closely, which required adjusting the weight of the model appropriately for each test condition. The density of aircraft in the 1930s was still quite low compared with the much larger values of later times. However, it was found that models for testing at the time of the opening of the Free Spinning Tunnel could usually be made to these requirements, when built mainly of balsa wood with the scaling of mass and moments of inertia and the correct locating of the centre of gravity being made with plugs of lead ballast.

In a similar manner to the procedure for the rate of descent, equating the lift force to the centripetal acceleration in the spin, as shown for the prototype spin in Part 1, and incorporating the result of equation 3 gives

2 2 (Ω R)a/(Ω R)m = 1 (4) and since the radius R of the helical path is a linear dimension that like all others would vary directly with the scale factor n, the predicted rate of rotation is given by

Ω a / Ω m = 1/√n (5)

Scaling factors for these and other quantities of interest derived by similar means are shown in Table 1, where it is seen that by using the weight density ratio in the modelling, the most frequently used quantities are scaled by simple powers of the linear scale factor n, though others still require the inclusion of the air density ratio σ.

* The relative density σ is ρa / ρm

184

Journal of Aeronautical History Paper No. 2015/03

Table 1. Factors for scaling quantities from model spinning tests, full-scale aircraft / model (model scaled by ‘weight density ratio’)

Quantity Factor Length n Angle 1 Time √ n Rate of descent V √ n Rate of rotation 1/√ n Radius of path R n Helix angle  1  Aerodynamic Force F n  Aerodynamic moment M n 5 Moment of Inertia I n

3.6 Modelling moments of inertia

For full dynamic similarity, the model is required to have the moments of inertia about the three principal axes properly scaled relative to those of the aircraft by the factor given in Table 1, and must have the centre of gravity located in the correct relative position in the airframe. It was expected at first that values of moment of inertia could not be obtained by calculation with an error of less than about 10% at best. To check this, the inertia of a full-size Bristol Fighter was measured (13). As had been reported also in the lecture by Scott-Hall, this was done by calculation from measurements of the periods of small oscillations when the aircraft was suspended as shown in Figure 9. For finding the moments of inertia about the longitudinal and lateral axes the aircraft was suspended as a compound pendulum, so as to produce a rocking rotation about the axis concerned. For a rotational oscillation about the normal axis, the model was hung from a pair of wires forming a bifilar suspension in torsion. The standard symbols employed for the recording of the moments of inertia about the longitudinal, lateral and normal axes were A, Figure 9 Suspension of full-scale Bristol Fighter B and C respectively. for measurement of moments of inertia (Reference 13)

185

Journal of Aeronautical History Paper No. 2015/03

It was found at first that there was a serious lack of agreement with calculated values, the measured ones coming out about 30 % greater. Then it was realised that the periods of oscillation would be affected by the inertia of the surrounding air which was also set in motion by the moving aircraft. Known as the ‘virtual moments’ for each of the three axes, these quantities were estimated by reference to supplementary oscillation tests with models. In these, the change in period was noted when a second measurement was made in air of one- fifth sea-level density in the RAE Altitude Chamber. The effect of the air could then be evaluated and a correction made for it in the full-scale tests. Later, the second measurement was made more conveniently in a small chamber containing hydrogen at ambient pressure. The necessity of determining the virtual moments due to the surrounding air was clarified by careful evaluation with a large aircraft-like model, made up of simple slab components of which the inertia could be calculated accurately. For this case the great effect of the surrounding air was confirmed, to the extent that if it had been neglected, the swinging measurements alone would have been ‘practically worthless’.

With measurements on an aircraft at full scale and model tests to determine the virtual moments, it was believed that the moments of inertia could be obtained with an overall error of about 5%. But it was duly concluded that this method was too cumbersome and lengthy for routine use. It was thought to be better to concentrate on ways of improving the accuracy of figures obtained by calculation, though this would mean accepting the tedium of having to work with the mass and location of the parts of the aircraft and its systems in quite fine detail.

No attempt had been made to formulate a theory for the associated air movement on the motion of the aircraft in these swinging tests. But the inertia of the surrounding air became a consideration in other situations in which there were accelerations, for example in flutter, which was already being addressed in theoretical work. However, once the spin had become steady (that is, unchanging with time) changes in the momentum of the air passing around the aircraft were represented fully in the aerodynamic forces and couples applied to it.

3.7 Spin recovery theory

A short R&M at this time showed however that efforts to extend the theoretical analysis of the spin to the unsteady stages were continuing (14). The authors were Bryant and Miss Jones (as still so designated at NPL). They aimed to produce an account of events at the beginning of spin recovery, in which various control actions were represented. Starting with the simple orientation of an aircraft in the prototype spin they considered the behaviour during the first few moments of the recovery, while it might be reasonable to suppose that principal quantities, such as the bulk motions of rotation and descent in the spin, would not yet have changed substantially. This required the inclusion of an acceleration term in each of the moment equations. These were then solved incrementally, as the aircraft was assumed to move through a series of steps very close together in time. The aerodynamic and inertial moments were recalculated for each step, using rotating balance data for the former.

Calculations were made of these initial stages of recovery for the two stable spin states o reported from full-scale tests with the Bristol Fighter – a steep spin with incidence 32.4 and

186

Journal of Aeronautical History Paper No. 2015/03

o a flatter one at 53 . With the simple aids to computation available at the time, the solution procedure, with very small time steps, was unavoidably slow and tedious, but fortunately from the results some clear conclusions could be reached. These focused on the initial changes to the attitude of the aircraft following a movement of the controls, namely that: a) an action which produces a rolling couple at this time will generally cause a change of sideslip, by which it is counteracted; b) an action which produces a yawing couple at this time, if moderate, produces little change of sideslip and is not counteracted; c) thus it is confirmed that the rudder is by far the most effective control to start an early recovery from a spin; d) ailerons assist or retard the start of recovery by virtue of the yawing couple that they produce and not by their primary function of applying a rolling moment; and e) reversal of elevators without other control movements may lead subsequently to the development of the faster, flatter spin from which recovery was likely to be more difficult.

Thus, the practical conclusion of this study was that in spin recovery it was best first to apply anti-spin rudder and not to move the stick until the rate of spin was observed to be falling off. These results did not run counter to prevailing ideas and practice, but the fact that those could now be underpinned by theory was potentially a valuable step.

4 Air flow in the spin 4.1 Wool tufts

It was seen that how the air moved around a spinning aircraft was an area of study that had been largely neglected. Without a clear impression of this flow it would be hard to understand how the aerodynamic couples that controlled the process were actually generated. This had been hampered by the absence of any proven means for investigating it, though one possibility for flow visualisation had been suggested by Flt Lt Haslam in 1928, as reported in Part 1. This was the use of wool tufts, which had appeared in the present context in the model testing of the Hawker Hornbill outlined above. Haslam had continued work with this method with super- vision by Professor Melvill Jones at Cambridge, in an investigation of stalls and spins with aircraft of the University Air Squadron, and this was reported in 1932 (15).

Haslam had flown an Atlas Mark I aircraft with and without tip slots, in stalled glides and spins and a Bristol Fighter, having interconnected slots and ailerons, in stalled glides only. Wool tufts were attached to the upper surfaces of the lower wings and on the tailplane. Ciné- camera records clearly showed the progressive nature of the stall as the regions of separated flow spread across the wing, and the presence of a standing vortex over the rear part of the wing before the flow became completely disorganised when the stall had fully developed. The local effect of slots in delaying the stall was shown clearly. Filming confirmed that the tailplane was also stalled at the higher incidences of both the stall and the spin. Thus was the utility of a simple technique for flow visualisation well demonstrated, and it continued to find uses at full scale and in wind tunnel tests thereafter.

187

Journal of Aeronautical History Paper No. 2015/03

4.2 Separated flows

Other impressions of the flow in the spin were reported at this time by B V Korvin - Kroukovsky (K-K hereafter) in an article in Aircraft Engineering on recent American work (16). Although outside the scope of the present paper of developments in the UK, this was referenced in several British papers after appearing in that widely-read journal.

One result for comparison with British work concerned the range of incidence over which spontaneous autorotation of wings could occur. Glauert’s first studies had shown that if the lift and drag characteristics were similar to those measured in ordinary wind-tunnel tests, this range would be very narrow for bare wings. That had been confirmed in tests of autorotation about a fixed central axis. But it had later become apparent that the behaviour of wings on complete aircraft could allow autorotation to take place from immediately after the stall to o incidences of up to 90 . This was most likely to occur in biplanes, particularly with wings having zero stagger and a low gap. K-K now reported that the narrow range of incidence for autorotation for monoplane wings was found only in situations in which there was no yaw. o As little as 5 of yaw removed the limitation, allowing autorotation to occur up to the high incidences experienced in the flat spin. This too was consistent with British work, that had shown how yaw produced large rolling moments at all incidences. As this was now traced to differences in the pressure distributions over the parts of the wing that led or trailed in the rotation, it was indicating that the lift and drag characteristics were unlikely to remain as measured in ordinary tests.

So far, there had been no attempt to represent these effects in theory. But K-K went on to consider the opposing effects of the coupling between yaw and roll, according to whether the yaw was positive or negative relative to the flight path. Because of differences in the relative magnitudes of the yaw and roll moments in the two cases, different terms in the moment equations could be neglected, giving two approximate expressions, the ratio of which could be interpreted to have practical consequences.

On the basis of these observations, a single Figure of Merit K had been devised, which was shown to discriminate clearly between aircraft that would recover readily from the spin (when K ~ 1.5) or fail to recover (K<1). By including results from both American and British aircraft, values for this figure were given for 12 that had recovered and 10 that had not. However, part of the argument for this procedure depended on an understanding of the likely modification of the moments that could be applied by the empennage and rear fuselage, which as yet was very uncertain, particularly when these lay within the wake of the wing,.

In illustration of this work, K-K included results of pitot-traverse measurements made by a Swedish researcher, of the flow behind stationary stalled wings in a wind tunnel at Göttingen (from 1925 the Kaiser-Wilhelm Institute for Fluid Dynamics)(17). Two examples reproduced in Figure 10 show the extent of the wake behind a rectangular wing of aspect ratio 6. At high values of incidence, it was found that the wake could diverge rapidly to many times the width of the wing chord. It was to be expected from this that the wake of the lower wing of a biplane would interfere significantly with the flow over the upper one. Although the spread was less at the lower incidence typical of the steep spin, the wake from the lower wing of a biplane

188

Journal of Aeronautical History Paper No. 2015/03 could then lie low enough to blanket some or most of the fin and rudder and perhaps the tailplane. K-K expressed justifiable surprise that, in view of the importance of this area to flight safety, these were the only wake measurements for separated flow that he could find anywhere in the literature at that time.

This study was noted by British workers, but the approach to defining the figure of merit K for the likelihood of spin recovery does not seem to have been particularly influential with them. This might have been because K required subjective values for empirical constants in the expressions for moments from the empennage, though at this stage that would seem to have been Figure 10 Spread of wakes behind a deeply inevitable. Nevertheless, the reported lack stalled wing (after Reference 17) of data on wake flows should perhaps have stimulated some useful work, for example in conjunction with rotating balance tests, which were now highly developed. These had already shown that moments when in rotation were very different from those measured in static conditions, no doubt to be expected when the airflow patterns themselves would be very different. Good facilities were in place, which could perhaps have been adapted to enable the wake flows to be mapped with probes rotating with the model. That was a topic which might have been taken up by a university, but the opportunity was seemingly missed, as nothing of this nature was reported in the UK in the period under review here.

5. Advice to the industry, 1933/34 5.1 R&M 1535 5.1.1 Presenting the case.

Irving’s report in response to the request for advice from the SBAC noted above was issued (18) as an R&M in 1933 . Like R&M 1001 in the 1920s, this report was a significant record, based upon a valuable summary of a wide range of recent work. The stated intention was again to assist designers, as far as possible given the state of knowledge at the time. It also carried the endorsement of the ARC, whose Spinning Panel had discussed it in detail and approved it. However, the hopes of the SBAC would not be fully realised, as the Panel concluded that ‘knowledge was not yet sufficiently complete to warrant the laying down of quantitative rules in simple definite form’.

It is not recorded that the SBAC had mentioned any specific views from its members arising from previous reports. Outlining the approach this time, Irving began by stating that it would not be concerned with how to avoid the involuntary spin or to recover from it, about which

189

Journal of Aeronautical History Paper No. 2015/03 the advice was now considered to be sufficiently good, but principally with preventing such a spin from developing into a flat spin. It was indicated that pilots had usually been able to o recover an aircraft from slow steep spins, but if the incidence rose above say 50 , recovery might become impossible because blanketing in the wakes from various parts of the airframe could seriously impair the effectiveness of the controls. Though shorter than R&M 1001, Irving’s report is quite detailed, so in summarising it here only the leading features are reviewed, mainly to draw out the continuing thread of advancement in the understanding of the spin and its representation to potential users.

In an introduction, a review is given of the orientation and motion of an aircraft in a spin and of the principal factors that were known to be involved in determining these. Irving’s intention was to provide a simplified account, which, although concentrating on the physical behaviour, would entail ‘some sacrifice of rigidity being made in the interest of clarity’. As in the case of Gates and Bryant in R&M 1001, his approach would include the representation of dynamics in terms of centrifugal forces and inertia couples, which was widely used by designers at the time. It will be convenient to summarise this work, especially the illustrations provided, to assist in the presentation of subsequent developments.

The diagram showing the applied forces in the spin is reproduced in Figure 11. The left-hand part of this again shows their general effects to be that the drag component of the resultant aerodynamic force on the aircraft balances the vertical force of its weight due to gravity, and the lift component balances the horizontal centrifugal force due to the rotational component of the path about the vertical spin axis. It is accepted, as first observed by Glauert, that the resultant force, the combination of lift and drag, always acted close to a direction perpendicular to the chord of the wing. Then it could be seen that in a flatter spin (with a greater incidence of the aircraft to its path), with the lift component becoming lower, the radius of the path becomes smaller and the rate of spin faster. However, it is not the attainment of equilibrium of forces Figure 11 Representation of an aircraft in a spin alone that imposes limitations on the (Reference 18) spin; these arise from the balance of couples, which determine the rate of rotation and the radius of the path on which the aircraft could settle into a stable condition

In the approach used by Irving, this balance was represented as one between couples generated by the aerodynamic forces and ‘inertial couples’ arising from the moments of inertia of the

190

Journal of Aeronautical History Paper No. 2015/03 aircraft in rotation. At this time, the distinction between a couple and its moment about a given point had been almost lost and the terms were used indiscriminately. This arose because it had been usual to represent the overall effects of an aerodynamic force applied at any part of the aircraft by two components. The first was as if an equivalent force had been applied at the centre of gravity, and the second action that of a couple equal to and of the same sign as the moment of the actual force about the centre of gravity. It could be assumed that this was understood, as otherwise it would be necessary to state the conventions of the subject at every use. These actions were generally resolved about the three principal axes of the aircraft – the longitudinal, lateral and normal axes. When the state was stable (unchanging with time) they could however be expressed without complexity about any chosen axis.

Irving went straight to the critical effect of yaw in the spin, the area in which the greatest progress had been made since the last major review. If there is no yaw, the longitudinal axis of the aircraft faces along its path, that is, the helical path formed by the combination of the vertical descent and the rotation in the spin. When the aircraft is yawed, there is a component of the airflow approaching its side from the direction opposite to that of the yaw, that is, it is side-slipping. Then there will be a nett sideways aerodynamic force acting on the side area of the aircraft, principally on the fin, rudder and rear fuselage. This force has three effects. The first is to act directly with or against the spin, according to the direction of sideslip. But there is also an indirect effect, due to the resultant of this lateral force acting somewhat above the fuselage axis and exerting a moment in roll. This produces a tendency to tilt the wings, in a direction which either raises the leading tip and lowers the following one, or vice versa. It was well known from rotating balance tests that tilting the wing limits the range of incidence at which it can autorotate. A third consequence of yaw that had been noted was that when a stalled wing is side-slipping, there is a powerful tendency to increase the lift near the wing tip towards which the slip occurs and to reduce it near the other wing tip. This affects the moment that the wing applies about the spin axis, the main means by which autorotation is maintained.

At high angles of incidence, the axis of the body is moved towards the horizontal and the forces on the tail and rear fuselage have a bigger leverage about the spin axis. By this illustration, Irving pointed out through the three mechanisms described that a ‘sufficient effective fin and rudder area, or its equivalent in depth of body near the tail, will prevent the flat spin from occurring’.

The manner in which the sign of the yawing displacement affects the rotation is then visualised by reference to the view on the right in Figure 11. Note that the aircraft is shown as seen from below. This is looking in the direction in which the anticlockwise yawing displacement being shown about the normal axis would by convention be positive. The aircraft is side- slipping and the spanwise orientation of the wings is changed. As the spin is right-handed in this illustration, the port wing is leading, the sideslip is towards that and the tip is lowered. So here the additional rolling moment is applied in the direction of the spin, and this would be a case referred to as one of ‘pro-spin’ sideslip, the opposite of that desired.

The position of no sideslip is when the aircraft is following the helical path directly. As that path is inclined to the vertical by the angle of the helix, the following tip is then already down,

191

Journal of Aeronautical History Paper No. 2015/03 at the same angle to the horizontal. To produce side-forces and moments that would oppose the rotation, the sideslip would have to be towards the following wing, requiring a yawing displacement which by convention would be negative. Here, Irving’s reading of events agreed with that of Korvin-Kroukovsky mentioned above.

If a spin is steady, there can be no nett acceleration of the aircraft about the spin axis and hence no nett couple acting on it. But Irving went on to represent how an aerodynamic moment is required about each body axis individually to balance a corresponding ‘inertia couple’ generated about that axis through its continually changing direction.

5.1.2 The inertia couples

The inertia couples arise from the distribution of mass in the aircraft, which can be expressed in terms of the moments of inertia about its three principal axes. They feature in the gyro- dynamic balances expressed in the Euler equations, which also include the rates of rotation about these axes, which are the components of the rotation in spin determined by the orientation of the aircraft to the vertical. It will be recalled from Part 1 that for the case of an aircraft in steady rotation, with components about the longitudinal, lateral and normal axes, equating the inertia and aerodynamic moments for each axis allows the equations to be cast in the reduced form

L = (C – B) q r , M = (A – C) p r and N = (B – A) p q (6)

These give the aerodynamic couples L, M and N required to be applied about the longitudinal, lateral and normal axes to maintain the rotations around them with component angular velocities of p, q and r in roll, pitch and yaw respectively. The component moments of inertia of the aircraft about these axes are represented by A, B and C respectively.

Up to that time designers in the industry had little reason to consider gyrodynamic effects in their regular work. Perhaps the only relevant situation was when calculating the couple required to be applied at the engine bearers when the direction of the axis of the rotating engine and airscrew was changing during manoeuvres. This could be done by a simple rule. Perhaps with this in mind, Irving again employed the approach based on ‘inertial couples’, the form in rotational motion corresponding to the use of inertial forces in linear motion, which was in general use then. The situation is presented as if the aerodynamic couple applied around an axis is required to balance an inertial couple, which is considered to have been caused by the rotation. This conceptual couple takes a form equivalent to the corresponding right-hand side of the appropriate equation in equations 6, but reversed in sign. In the report, this supposed state of ‘dynamic equilibrium’ between aerodynamic and inertial couples was formulated by balancing them around the lateral (pitching) and the normal (yawing) axes of the aircraft.

5.1.3 The situation in pitch

The couple in the pitching plane is represented by the diagram reproduced as Figure 12. The centrifugal forces, distributed along the inclined length of the aircraft, produce an inertial

192

Journal of Aeronautical History Paper No. 2015/03 couple about the lateral axis, as previously explained by Bryant and Gates. The distrib- uted mass is represented here as equivalent to two concentrated masses located on opposite sides of the centre of gravity. The implication is then that the rotation itself has generated an inertial couple due to centrifugal forces on these equivalent masses. Its direction is such that it would ‘tend to increase the incidence’, and that would have to be opposed by an aerodynamic couple to produce the state of dynamic equilibrium. The moment of this couple is represented here as if being applied by a spring (which would need to have non- Figure 12 Origin of 'inertia couple' in pitch linear properties to correspond with the (Reference 18) complexity of the aerodynamics). Its sense would be anticlockwise in the view of Figure 12, which would be negative in the standard definition for rotation about the lateral axis, which points into the page in the direction of the starboard wing.

Although endorsed by the ARC Spinning Panel, there is a risk with this approach of implying that the relevant effects of rotational inertia about the lateral axis arise solely from mass within the pitching plane. The two separated concentrated masses would be chosen and placed so as to give rise to the moment of inertia B for this axis. However, the Euler equation for the required couple in equations 6 shows the moment of inertia part to be the difference (A - C) between the moments about the other two axes. When this issue arose in the earlier study by Gates and Bryant, they had pointed out that the numerical values of the moments of inertia about the principal axes of an aircraft satisfy the identity C = (A + B) quite closely, so that B is equal to (C - A) to a good approximation. Irving simply states without explanation that ‘the magnitude of the inertia couple in pitch depends on the value of the difference, C – A, of the moments of inertia about the transverse [i.e. normal] and longitudinal axes’. With the change of sign due to use of the dynamic equilibrium concept, this is in accord with the corresponding term in equations 6. Having been given an explanation of the situation in pitch that involved only masses that contribute to the moment of inertia about the lateral axis, it would be confusing for a reader coming to the subject for the first time to learn that somehow it really depended on the values for two other axes that each lay at right angles to that.

The aerodynamic couple which is effective in the pitching plane arises mainly from the moment of forces produced by the flow over the tailplane and elevators. It was claimed by Irving that this moment could be estimated to a sufficient accuracy from an ordinary wind tunnel test of a model at high incidence but without rotation. The presentation of the model to the airstream would presumably have to include an angle of yaw that represented any side- slipping for a particular case.

Irving went on to illustrate the main trends in the effects of design values by reference to calculated results, using the dimensionless group Ω s/V as a basic characteristic of the spin,

193

Journal of Aeronautical History Paper No. 2015/03 representing the ratio of the contributions to the relative wind of the rate of rotation Ω and the vertical rate of descent V. Values of this group had been obtained from full scale measurements and were expected to lie generally between 0.5 and 1.0. In the calculated results, a further parameter is the angle of incidence α of the wings to the path of the centre of gravity. He also introduced the method of assigning magnitudes to applied moments and couples involved in spinning, which was subsequently adopted by others for a time. In this, the moments are first 2 made dimensionless by dividing by the group ρ V Ss. This group is similar to that already used for making the moment of a wing section about its aerodynamic centre dimensionless, except that the linear dimension chosen for this latter case is normally the mean chord. Irving used the semi-span s, as dimensionless groups having this denominator arise in theoretical 3 analysis of the spin. He also multiplied the result by 10 so as typically to have values within the range 1 to 100. This scale was chosen to provide ease of comparison with other standard dimensionless moments; for example the moment from full rudder deflection in level flight, which would have been generally of order 10 in these units at that time. (For that, the linear dimension used was the distance of the aerodynamic centre of the rudder to the centre of gravity of the aircraft, which is usually of the same order as the semi-span). Where values of aerodynamic moments are quoted hereafter, they will be in dimensionless units as first introduced to the spinning literature by Irving.

Little in the way of advice to designers was offered in this section of the report, beyond that a large value of the moment of inertia difference (C – A) is undesirable. This is equivalent to B, as it appears in the inertial couple which is represented as tending to increase the incidence, as shown in Figure 12. For single-engine types, the largest concentrated mass contributing to B is the engine, which for reasons of balance is already located not far from the centre of gravity, so there would be little room for designers to vary this. The position with multi-engine types would be more debatable, but at this time the spinning characteristics of these had scarcely been considered.

5.1.4 The situation in yaw

Irving then illustrated the position for rotation about the normal axis (Figure 13), representing the aircraft when it is spinning with sideslip due to a displacement in yaw. The aspect is the same as in Figure 11, the view being from below the aircraft, as seen from outside the spin.

In the view on the left of the Figure, the mass in the aircraft was shown as if concentrated at two points in the fuselage on either side of the cg, as discussed before. This represented a case in which the moment of inertia B about the lateral axis predominates. An inertial couple arises about the normal axis due to centrifugal forces on these masses, acting in the sense of tending to increase the yaw. If on the other hand the mass had been concentrated in the wings, as in the view on the right of the Figure, the moment of inertia A about the longitudinal axis predominates, and the inertial couple would act to decrease the yaw. Irving defined the position to be ‘stable’ or ‘unstable’ according to the sign of the difference between the moments of inertia (A – B). He then embarked on a lengthy exposition (continued in an appendix to the report) in which he considered how this factor might bear on a tendency for a spin to become flat.

194

Journal of Aeronautical History Paper No. 2015/03

The tendency for a spin to flatten would depend also on the extent to which the aircraft can generate autorotation, as outlined earlier. Wind-tunnel work had shown that for a monoplane, spontaneous autorotation was confined to a narrow range of incidence immed- iately following the critical value at which the stall begins. This was also a characteristic of biplanes if the two Figure 13 Origin of 'inertia couple' in yaw wings were mounted with moderate (Reference 18) values of the gap/chord ratio and the stagger. When the incidence was beyond this small range, the aircraft did not rotate spontan- eously and if rotation was taken into this range by some means it was quickly damped out. On the other hand, it had been found that autorotation of biplanes with small gap ratios and zero stagger could occur at any incidence beyond the stall, as a result of changes to their aerodynamic characteristics caused by interference between the flows over the lower and upper wings. For Irving’s interpretation, in the first case the wings are said to be ‘stable’ in autorotation, as the aircraft would not easily be driven to high incidence in the spin. In the second case it is ‘unstable’, as once a spin had developed there would be an autorotative couple that tended to increase the incidence.

There could thus be four possible combinations of ‘stability’ and ‘instability’, two for the inertial couples and two for the aerodynamic couples, according to the sign of the yawing displacement as discussed in 5.1.1. Unsurprisingly, progression to a flat spin was shown to be least likely in cases where the difference in moments of inertia (A – B) and of the aerodynamic characteristics were both described as ‘stable’. Other combinations were considered at some length, though the consequences tended to be hedged about with more qualifications and were not so sharply defined.

After reaching these conclusions, Irving ended this section by stating that, on the basis of his calculations, ‘even when the combination of rolling inertia properties is markedly unfavourable it requires no impractically large vertical surfaces aft of the centre of gravity, provided these are efficient in the spin, to counter the flat spinning tendency’.

5.1.5 Some practical illustrations

Irving illustrated his conclusions by showing values of the couple required from these surfaces, as estimated for a number of aircraft, including those for which the spinning behaviour had been studied at full scale. A total of eight cases could now be covered, by including the tests with the Bristol Fighter before and after modifications to the rear fuselage. The ‘vertical surfaces’ involved include the presented side area of the rear fuselage, and in a new present- ation, this contribution of the body was expressed through a characteristic geometrical factor F of the fuselage to be known as its ‘body damping ratio’. This is similar to the tail volume coefficient already used in considerations of lateral stability. It is given by

195

Journal of Aeronautical History Paper No. 2015/03

2 2 F = ∑(dAf x ) / S s (7) which expresses the sum of contributions to the restoring moment when the aircraft is yawed from elements of the presented side area of the fuselage dAf (this element has the symbol dS” in Irving's report). If dAf is located at a distance x from the centre of gravity, its displacement in yaw is proportional to x and the moment arm of the aerodynamic force introduces a further 2 multiplier of x. Division by S s is to make F non-dimensional. The idea is to give a rough numerical value to the effect of the area, taking into account its position relative to the centre of gravity. For this purpose, only areas aft of the centre of gravity were to be included. Though not specifically mentioned, for aircraft with engines in the nose, the comparatively small proportion of side area forward of the centre of gravity would not have a powerful contrary effect.

Some model results obtained by Irving are shown in Figure14 for the particular spin state o with Ω s/V = 0.7 and α = 60 . This gives calculated values of the dimensionless yawing moment with sideslip versus the body damping ratio F. Although the Figure does not establish a unique relationship, the results for different aircraft lie between two limiting lines, according to the shape of the cross-section of the fuselage. Those with circular sections would require about twice the value of F to balance a given inertial moment than ones with flat-sided sections. Here, Irving was drawing on early results of a study, not yet published, that will be reviewed later. At this point, the presentation in Figure 14 might be noted as an early example of visual forms that would arise later in Section 8, in which boundaries would be placed on diagrams to separate cases of satisfactory from unsatisfactory performance in the spin.

Figure 14 Effect of lateral area of body on yawing moment in spin (Reference 18)

196

Journal of Aeronautical History Paper No. 2015/03

In any interpretation of these results, it would be essential not to overlook the qualification given by Irving that in making calculations, it had been necessary to assume that the surfaces would be ‘efficient in the spin’. In practice this would be when they were not significantly blanketed by wakes from other parts of the airframe. At this time there was still little that could be said on how the designer could be sure of this in a given case.

5.2 Compilation from R&Ms 1535 (18) and 1644 (19)

Irving’s report ends with a short collection of conclusions relating to design. Perhaps the most significant of these are the requirements as seen by him for ensuring the balance of couples in yaw. Published results for the magnitudes of yawing moments due to rolling are collected into tables. These give the contributions for wings, fuselage (without fin and rudder) and fin with rudder undeflected, together with some for rudder fully deflected, as during an attempt to recover. For typical ranges of the spin parameter Ω s/V and incidence α, the data are listed for various types of fighter aircraft and for the Fairey Type IIIF, with and without floats. This is a substantial collection of material, including results from both rotating balance and full- scale testing. However, the difficulty of being able to make use of these at the design stage, as pointed out by the Spinning Panel, can be seen from the wide variation shown. As a o ‘rough guide’, Irving gives the following values for fighter aircraft with α = 60 and Ω s/V = 0.7, representing a typical flat spin:

Contributions from parts of the airframe to yawing moments due to rotation in spin (Unit as given by Irving in sub-section 5.1.3) Wings +10 to - 10 Body - 5 to - 40 Fin & rudder +2 to - 30 Rudder against spin - 2 to - 10

Irving’s own suggestion was that a composite figure of - 45 units could be regarded as the lower limit required from the combined body, fin and rudder to ensure that the aircraft would remain in the steep spin and not switch to a flat spin.

R&M 1535 (18) was followed the next year by another on similar lines by Gates and Francis (19), which gave a sharper focus to the effects of inertia outlined by Irving, by showing some results of general application. The subjects now were several idealised types of aircraft, each with variation in characteristics such as the moments of inertia A and B and the position of the centre of gravity. Where it was reasonable to do so, certain quantities for the aircraft were assumed invariant, taking typical values according to the type. Aerodynamic couples in roll and yaw were obtained from rotating balance data for monoplanes and biplanes, together with flight data available for the Bristol Fighter, with and without slots. These were reduced to the common dimensionless form and the output from the calculations given in terms of the spin parameter and incidence as used by Irving.

197

Journal of Aeronautical History Paper No. 2015/03

Because of the extensive number of combinations to be covered, a total of 48 graphs are required to express the results. The prospect of having to extract guidance from so many figures would be daunting to a busy designer. But in any case, Gates and Francis acknowledged that there was little room in practice to improve spinning characteristics by adjusting the distribution of the moments of inertia, which would ‘usually give the designer a good deal of trouble’. This was apparent enough, given the conventional configurations towards which aircraft design had converged and the limited opportunities that this provided in practice for relocating components of significant mass. However, it would not be fair to conclude that studies such as these were without value. It could not have been foreseen, before the work was done, whether or not there would be strong indications that the distribution of mass was of major significance, and if so whether designers should now consider possible new configurations.

This is exemplified by the appearance in the wind-tunnel test programme of a model of the Westland-Hill Pterodactyl aircraft, of which the Mk II version is shown in Figure15 (20). As this was a tailless design, the wings were swept back to provide longitudinal stability. Pitch and roll control were by elevons, rotatable wing-tips which acted together as elevators and differentially as ailerons. Directional control was by small rudders mounted beneath the wings. There being no tailplane, adjustment for trim was achieved by varying the wing o sweep through a range of about 5 . As the fuselage was relatively much shorter than in conventional types, the moment of inertia in pitch B was significantly reduced compared with a typical value for other aircraft of similar span, while that in roll A, dependent largely on the mass distribution in the wings, was not greatly different. The dimensionless factor containing (A – B) was then about 0.4, a high positive value which by Irving’s definition would represent a very ‘stable’ inertial characteristic. (For typical single- engine aircraft of conventional layout, this factor was rarely much above zero and had been found as far towards the ‘unstable’ side as - 0.4 (18)). When a large positive value appeared in conjunction with a monoplane wing, which was expected to be in the ‘stable’ aerodynamic category, Irving’s theory would predict that the aircraft would have satisfactory spin and recovery Figure 15 The Westland-Hill ‘Pterodactyl II’ characteristics. (FAST Museum Collection)

Some limited testing of its spinning characteristics were made for the Pterodactyl IV at RAE, both in the vertical tunnel and at full scale, with good agreement between them (21). Spinning was difficult to induce unless the centre of gravity was far back, and then was of the slow, steep variety, with a large radius and helical angle. The rudders, now in the wake of the stalled

198

Journal of Aeronautical History Paper No. 2015/03 wings, were almost ineffective, and were enlarged for full-scale trials, but it was found that recovery could be obtained rapidly using the elevons alone. As the aircraft had no separate elevators, there had been concern that the forces on the elevons in the spin would strongly oppose the forward movement of the stick needed to complete the recovery. A device employing a pendulum weight at the outer wing-tip was fitted to mitigate this, and in no case was any difficulty in recovery experienced. To that extent, Irving’s prediction was verified.

Nevertheless, the overall conclusion of this work was that for limitation of the tendency to develop a flat spin ‘good body design is a more powerful safeguard than attention to good mass distribution’. This was to be found mainly in the shape of the rear fuselage in side elevation, through its contribution to the restoring motion set up when there was yaw. Irving also included here some results from an on-going study, which was published later. These illustrated the extent to which the cross-sectional profile of the rear fuselage was also important. This was because of its effects on the wake formed by the fuselage, which could envelop the empennage in the high incidence of a flat spin, directly reducing the effectiveness of the tailplane in pitch and that of the fin and rudder in yaw.

These accounts had brought out the new understanding of the spin, through appreciating the effect of yaw and sideslip on the tendency of a spin to progress to the fast flat type from which recovery was unlikely. Extensive consideration had been given to the aerodynamic and inertial characteristics of an aircraft that should promote an attitude in yaw which would discourage that progression. This was valuable as a record of progress in research, but it had not led to the emergence of the hoped-for simple design rules to ensure safety in the spin. Specific numerical values or defining expressions would be required to support those. Irving, Gates and Francis had shown the extent of advances in current understanding, but the terms were still not sufficiently precise. For the time being this would largely give designers further support for practical empirical measures, such as enlarging the size of the fin and rudder, which had been identified already in earlier papers published in professional journals and articles in popular technical magazines.

6 Further developments in spin testing 6.1 A coordinated programme

Work in the early part of the 1930s had shown how effectively the analytical procedures for representing the spin, begun by Glauert in 1917, had now matured. Throughout that time, the analytical and experimental sides of making progress in understanding the spin had informed each other at every stage. While this process would remain open, the scope of the theoretical side would continue to be restricted to the steady spin. It was becoming clear that a whole new field of unsteady aerodynamics would have to be addressed to make much further progress into representing the recovery phase. The more immediate developments would be on the experimental side. That was reflected in a significant shift of views at the ARC at this time, where it was reported that ‘The Spinning Panel advised that theoretical work could, with (22) advantage, be postponed in favour of the generation of more ad-hoc design data’

199

Journal of Aeronautical History Paper No. 2015/03

The advent of the vertical wind tunnel had evidently provided a strong new impetus to the study of spinning with models, including actions for recovery. But earlier methods were not made obsolete by that. A coordinated programme of work was now begun, in which a family of model aircraft was tested by RAE in the Free Spinning Tunnel, with systematic variation of one design feature at a time, while the aerodynamic couples due to body and empennage of various shapes continued to be measured with the rotating balance in horizontal tunnels at NPL. Studies on particular aircraft would be included when interesting or unusual spinning behaviour was reported from full-scale flight test programmes at RAE and A&AEE.

A new version of the NPL rotating balance had been devised, in which the radius of the rotating path could be offset from the axis, so as to reproduce the helical form of the path in the spin more closely (23). The maximum radius of the path that could be set was 12in, with rotation up to three revolutions per second. Shown in Figure 16, the balance now allowed measurements to be made of the couple due to yaw about the normal axis of the model and the rolling couple due to rotation about the spin axis. These were obtained from the deflections of stiff springs, which moved the iron cores of solenoids by small amounts, generating signals that were transmitted to recording instruments via slip-rings and brushes. Proposed developments would allow additional measurements to be made of the pitching and rolling couples about the body axes and of the total drag of the model.

Shortly after this, a further attempt was made to use the Whirling Arm at NPL in spinning research (24). Originally devised for testing airscrews, this was robust enough to carry fighter models of around 3ft span in a circular path of 30ft radius, at forwards speeds up to 100ft/s. At this scale the rates of sideslip would represent typical conditions in the spin. As reported in Part 1, this device had not provided consistent results in earlier trials, but it was brought back into action after modifications to limit the effects of swirl induced in the Figure 16 Revised NPL rotating balance with air within the building in which it Bristol Fighter model on offset mounting operated, thought to have been a (Reference 23) detrimental factor. The model was now mounted in a rigid octagonal frame, equipped with balances to enable the induced couples to be measured about the roll and yaw axes, as shown in Figure 17. Tests were first made with a model of the Hawker Hornbill, as shown, that revealed a large anti-spin rolling couple, beginning soon after the stalling angle had been exceeded. It was now thought likely that the strength and sign of this couple during sideslipping might account in part for the marked stability of the Hornbill at incidence above the stall and its ‘advantageous properties relative to spinning’. A series of measurements then followed, using wing models of various aerofoil

200

Journal of Aeronautical History Paper No. 2015/03 sections and tip shapes, with and without slots, together with pressure plotting and camera observations of the airflow over the upper surfaces, using wool tufts and smoke. These confirmed the presence of a general outward flow over the model, together with a strong eddy on the upper surface of the wing near the inner tip, of which there had been other indications previously. The flow visualisation and pressure plotting were found to be consistent. Some concerns remained about possible complications due to enhancement of the induced outflow by centrifugal action with this apparatus, so further modifications and tests were planned.

6.2 The monoplane configuration

There was said to have been an enduring antipathy at the Air Ministry towards monoplane Figure 17 Supporting frame of NPL aircraft, arising from the relative frequency of Whirling Arm, with 3ft span Hornbill model structural failures with examples of this layout (Reference 24) in the early days of manned flight. Subsequent developments in the general aviation sector had soon provided successful light aircraft of this configuration for transport, touring and participation in air racing. In the Schneider Trophy competitions, entries had much higher wing loading and were designed for maximum speed, characteristics likely to be required in fighter aircraft of the future. As the contests proceeded, the trend had been towards monoplane designs and in the later ones all the entries were of this form.

The biplane had been the norm for fighter aircraft in the RAF up to the start of the 1930s. Some fine biplane machines, well-adapted to the need for agility in aerial combat, had entered service and had also sold well to forces abroad. For this layout, the tendency had been to select thin wing sections that kept the weight of wings low and provided sufficient flexibility to allow minor adjustment to be made to rigging angles through tensioning the interplane bracing wires. But the choice of an aerofoil section for the wings of an aircraft for a given duty depends on many factors. Sections with greater thickness ratios had been investigated and showed some aerodynamic characteristics that would be more favourable for the operational requirements that were now emerging.

Air Staff planners had concluded that political events taking place in Europe would bring a new threat to the British Isles from concentrations of fast bombers approaching from the quarter between east and south. Effective defence in that situation would be possible only with interceptor fighters having a combination of the highest speed, rate of climb and concentrated fire-power that current technology could provide (25). As the warning of approaching enemy aircraft would be very short, the rate of climb would be a critical factor,

201

Journal of Aeronautical History Paper No. 2015/03 and a key aerodynamic characteristic for that would be a high lift-drag ratio. Somewhat thicker sections were better in this respect, and the greater depth available when these were used would enable the required wing strength and stiffness to be obtained in the cantilever monoplane configuration, assisted by the now-established trend to all-metal stressed-skin construction.

The requirements as set out by the Air Ministry took some time to become clear, leading to the issuing of a series of specifications with increasing demands that were a challenge to the industry. The balance of factors between biplane and monoplane layouts was about to shift decisively in favour of the latter, though this took effect over a period of some years. One technical element in the choice of wing thickness was that the more rounded leading edge shape of the deeper sections delayed the onset of separation to a somewhat higher incidence and the stall developed more gradually than with thinner wings. Possible implications of this were soon explored in investigations of the spinning characteristics of the monoplane layout.

An early example of that was a study addressing the basic question of where the single wing should be best placed in relation to the fuselage. With some suggestion of haste, the objective was, ‘to compare by as simple a method as possible’ the spinning properties of high and low wing monoplane configurations, which was addressed by measuring the rates of autorotation of models about an offset axis in the wind tunnel (26). For each test the model was mounted on a trunnion by which the incidence and yaw angles could be set within ranges appropriate to those occurring in the spin. This was mounted on a transverse arm, adjusted so that the radius of the offset path of the centre of gravity of the model about the spin axis was scaled roughly as indicated by theory for each setting of the angles employed. This arrangement is shown in the lower part of Figure 18.

It was found that with the wings in the low position the model autorotated considerably faster than when it was in the high position, giving rise to a speculation that a low-wing monoplane could be more likely to spin flat than one with a high wing. The opportunity was taken also to make the model tests with different positions of the tailplane, first in the common position on the fuselage centre-line and then on top of the fin and rudder. The difference in the results showed again that when the tailplane was in the lower position, even a large fin and rudder could be completely shielded by it in spins at high angles of incidence. These additional results, being in agreement with earlier ones, tended to give credence to the finding in relation to the most favourable location for the wing.

The monoplane layout was chosen when a modern illustration was required for a revised representation on the covers of R&Ms of the standard conventions to be used in analysis. The new form is shown in Appendix 1. The accompanying table of standard definitions and symbols showed the revisions to the ways of making leading quantities non-dimensional that had recently been agreed internationally. These fixed the linear dimensions to be used for various purposes (span b generally replacing semi-span s, for example). The most noticeable 2 change for European aerodynamicists was the use of the dynamic pressure ½ ρ V in the denominator of lift, drag and moment coefficients, with C as the standard symbol for these 2 coefficients in the new form (previously, the divisor had been ρ V with k the symbol for coefficient). This new convention still applies today. It has been used for coefficients from

202

Journal of Aeronautical History Paper No. 2015/03 the outset in the present paper to avoid having to make a change at this point. Uptake of the convention was not uniformly or promptly carried out, and the old conventions continued to appear occasionally, even in reports from the research establishments.

Modified form in autorotation tests at NPL (Reference 26) Figure 18 The Vickers Jockey monoplane

The coordinated programme on spinning behaviour continued to be aimed at providing data on the effects of design features for the guidance of the industry. Some additional tests were interleaved with this, to follow up the central issue of the agreement between the results of free- spinning model tests and experience with the corresponding aircraft at full scale. Information

203

Journal of Aeronautical History Paper No. 2015/03 continued to be obtained through specific flight research trials at Farnborough and from reports remitted from manufacturers' tests and from Martlesham Heath on behaviour during routine evaluation tests, with particular interest in the early monoplane designs that were now coming forward.

An example of these comparative studies concerned a low-wing monoplane, described in reports only as an ‘Interceptor Fighter’, but identifiable as the Vickers Type 151, known unofficially as the Jockey (Figure 18). This aircraft had been designed to the early specification F.20/27 for an ‘interception single-seater day fighter’, required to engage enemy bombers having a speed of 150mph at 20,000ft. It had a forward-looking layout, with the unusual feature of an engine mounting that was hinged so that it could be swung sideways for ease of maintenance, without requiring any services to be disconnected. However, it had been given a fuselage of basically circular cross-section (though with a corrugated skin) and a fin and rudder located directly above the tailplane, features that were already strongly indicative of a liability to develop a flat spin. Early flight tests indicated a lack of torsional rigidity of the rear fuselage, which was redesigned by Barnes Wallis, and other improvements made, including the fitting of a Townend ring to reduce the drag of its Bristol Mercury radial engine. A model of the aircraft in this form is shown in the lower part of Figure 18 as the subject of the autorotation tests made at NPL.

Although not chosen to proceed to full service trials, this aircraft had some advanced design features and had shown good stability near the stall, so the prototype was sent to A&AEE for flight tests. These began with a spin to the right from 15,600ft. It was recorded as steep at first, but unsteady, with the nose rising and falling during the rotation. When the stick was released the spin became faster and flatter. After about six turns, the rudder was reversed and the stick moved fully forward, in the usual moves for recovery, but the flat spin had become established and there was no response. Succeeding events followed a familiar pattern. The pilot reported trying every artifice at his disposal, including stopping and starting the engine with various combinations of control position, though producing no change. In these attempts, he lost count of the number of turns, but the aircraft had continued to spin downwards for about 10 000 feet. His laconic third-person record continued “At about 6,500ft, the engine was throttled back, the rudder held reversed and the control column pulled back, but without effect. During this period the elevator and aileron controls appeared to offer no resistance to movement. As the altitude was by this time little more than 5,000ft, the pilot decided to abandon the aircraft. Examination of the wreckage showed that the aircraft had struck the ground whilst spinning in a flat attitude”.

Model tests were now made in the vertical tunnel at RAE for comparison with those at full (27) o scale . It was found in these that with elevators up, the spin was steep (inclination 20 – 30 to the vertical) and recovery was quite satisfactory (about 2.5 second full-scale equivalent after o moving the controls). But with elevators down the spin was flatter (45 – 60 ) and the recovery much slower (13.5 – 16.5 second at full scale). It would be of little significance that the model had been recovered from this condition, as in practice a pilot would rarely want to wait that long to see if the controls had any effect, and would probably have gone on to try some different actions. It had also been found that in the steep spin with the elevators initially held in the up position, if these were merely released it immediately became flat.

204

Journal of Aeronautical History Paper No. 2015/03

It was considered that the model and full-scale results were telling essentially the same story. In both cases, the slow steep spin and recovery had been normal, though the accompanying nutation reported at full scale suggested that the aircraft had been close to instability. The fast flat spin had followed in both cases when the elevators were released, and as this was also the result with the model when they were held down, it was concluded that when released in the steep spin, the aerodynamic forces on the elevators had driven them into the fully down position, and the flat spin had then followed. The pilot’s observation that in the flat spin, there had been “no resistance” to movement of the controls was clearly a consequence of the control surfaces all being blanketed in the wakes of the stalled wings and tailplane at high incidence. Once this had occurred, the pilot would be left with no option but to abandon the aircraft.

Accumulated experience could suggest modifications that might reduce the tendency of a machine to move into a flat spin, and the vertical tunnel now provided a means by which these could be checked promptly. As an example of what became a typical action when spin troubles emerged, the modifications made to the models of the ‘Interceptor Fighter’ are outlined in Figure 19 and the results shown in Table 2 below. In the tests, in all cases the time per turn lay between 1.5 and 1.8 second equivalent at full scale, and the time to recovery with elevators down is given in the Table. For some of these tests, a check was made of the effects of increasing the lateral moment of inertia B by 20 % and of moving the centre of gravity forward. These were factors that were known to make the flat spin more likely, and their use represented the first attempts at determining whether there had been any margin of safety for the model test results in respect of potential future changes in those quantities.

Table 2 Spinning tunnel tests of ‘Interceptor Fighter’ with modifications to rear fuselage and empennage (times of recovery after movement of controls are the full-scale equivalents) (Reference 27)

Mod Time of Recovery, Modifications no seconds (elevators down) Narrow aft end of fuselage to elliptical section and 1 [flat, 10.3 – 12.2] blend to vertical stern post 2 Cut away inner edges of elevators [no further improvement] Raise tailplane to top of fuselage; [steep, 4.2 - 4.7] 3 with increased B and centre of gravity forward [variable incidence, 7.0 – 16.2] Incorporate stern fairing into rudder, reducing height 4 [variable incidence, 4.2 - 8.4] of fin/rudder to maintain same side area Move tailplane forward to reduce screening of rudder; [steep, 5.2 – 5.4] 5 with increased B and centre of gravity forward [steep, 7.3 – 8.9] Return tailplane to original position fore-and-aft but [steep, recovery by elevators 6 raise part-way up fin and add dorsal extension alone] Retain circular fuselage but place tailplane entirely behind fin and rudder. Reduce chord of tailplane to 7 [steep, < 5.0] give same tail volume as before. Fin and rudder volume unchanged but hinge line canted forwards

205

Journal of Aeronautical History Paper No. 2015/03

Figure 19 Modifications made to the 'Interceptor Fighter' model (Vickers Type 151 Jockey) (Reference 27)

The possibility of obtaining results such as these so promptly shows the great advance in spin research afforded by the availability of the vertical spinning tunnel, particularly the opportunity to make direct checks of the effects of potential changes to a design. As the results were also published, interested designers could note what features were most effective. In this case it had been shown that only relatively drastic modification could eradicate the tendency of the model to spin flat.

6.3 Safety and spin testing

For a new type, designed to an Air Ministry Specification, with hope that it would be selected for entry into service, spinning behaviour would first be examined in tests by the contractor. If that and other characteristics were found to be satisfactory, a prototype might be sent on to A&AEE at Martlesham Heath (or MAEE at Felixstowe for naval aircraft) for acceptance trials.

206

Journal of Aeronautical History Paper No. 2015/03

Clearly, there were risks at any of these stages, but there was most concern about cases which had suddenly shown dangerous tendencies at some point, while not having been troublesome in earlier stages. An example was the Bristol Type 133 monoplane, designed to the day-and- night fighter specification F.7/30. One of the first aircraft to use Alclad sheet aluminium, this had a gull-wing layout, with main wheels retracting into pods at the lowest points, as shown in the upper part of Figure 20. The contractor’s flight trials had found its spin to be of the regular steep variety, presenting no difficulty with recovery. But with the trials completed, one of the Bristol test pilots T W Campbell, when taking the prototype up for handling experience, entered a spin with the undercarriage inadvertently left down. A flat spin rapidly developed, from which recovery was impossible. Campbell abandoned it at 2,000ft to land safely by parachute and the aircraft was destroyed. Seen in the lower part of Figure 20, the wreckage shows the characteristic appearance recognisable as the outcome of a flat spin, where the aircraft is close to horizontal and has no forward speed at the point of impact.

Figure 20 Bristol Type 133 monoplane fighter to Specification F.7/30.

Top : Sole prototype, showing retractable undercarriage.

Bottom : Characteristic form of wreckage after flat spin (National Aerospace Library Collection)

207

Journal of Aeronautical History Paper No. 2015/03

Careful attempts subsequently had failed to reproduce this spin in model tests. Particular attention was given to the effect of extending the undercarriage, as this would change the moments of inertia A and B, both by magnitude and in their proportions relative to each other, by moving a significant mass further from the centre of gravity. But in subsequent model tests no differences could be found with the undercarriage extended or not. This and other cases caused RAE to revisit all aspects of spin testing.

One immediate matter of concern from the safety standpoint was the potential effect of the altitude on spin behaviour. In the vertical spinning tunnel, it became usual to make tests representing conditions at two operating altitudes, generally around 5,000ft and 15,000ft. In Section 3.5 it was shown that for this the model should be ballasted so that it had the same ‘weight-density ratio’ (W/ ρ Ss) in the tunnel conditions as that of the aircraft at the relevant operating altitudes. Results so far had always shown that recovery was easier at the lower altitude, and that had agreed with experience in routine full-scale spin testing at A&AEE. A study was now made by Gates and Stephens to see if any clear implications for aircraft design might arise from this (28). Involving both theory and experiments with models, that study was to lead to a development in the procedure for model testing which would have long-term significance.

The theoretical analysis confirmed that the yawing moment required to maintain a steady spin tended to be less at a lower altitude. As seen earlier, this moment is provided mainly by the side-areas of the fin and rear fuselage exposed to air flow due to sideslip. Using relevant properties of aircraft in typical cases it was found on this basis that the spinning qualities would always be better at lower altitudes, as had been observed. A median example showed that the yawing moment at 15,000ft would need to rise by something of the order of 6 units (anti-spin) to hold the same incidence as at 5,000ft (the units for moment used here and below are those introduced by Irving, as reported in Section 5.1c). It would normally follow that recovery would be easier at the lower altitude, assuming that the moments from the side area and from the rudder when activated had not been compromised by shielding or other wake effects.

However, in further analysis it was also found that there could be an unexpectedly large adverse effect of altitude on the rolling moment due to sideslip, which had hitherto been thought to be comparatively unimportant. This was dependent upon a more obscure complex of aerodynamic and inertial characteristics, and the range of possible values for the effects on the required yawing moment was quite wide. But in conceivable circumstances, this term could reverse the sign of the change in required moment, meaning that recovery would become more difficult as the altitude decreased. In a practical case, that would clearly be a very dangerous trend. As well as being a hazard for pilots in general, it would have implications for the safety of aircrew engaged in spin test work.

Some results were available from spin testing in the vertical tunnel which Gates and Stephens were able to bring to bear on this problem. These were from a model of an un-named single– seat biplane, for which the only details given were that the wings were highly staggered, and that the lateral moment of inertia B was greater than the longitudinal value A. When its characteristics were reproduced with a weight-density ratio equivalent to an altitude of 4,000ft,

208

Journal of Aeronautical History Paper No. 2015/03 spin recovery had seemed very satisfactory, and an idea emerged for determining what margin of safety there had been in that situation. This was assessed by applying a series of increasingly adverse yawing moments (i.e. in the pro-spin direction) by means of a small vane attached to the model's inner wing tip. The moments applied at different vane settings were calculated and converted to full-scale equivalent and the corresponding time taken to recover from the spin measured for each setting. Then the tests were repeated with the model scaled so as to represent conditions in a spin at 15,000ft.

The results for two sets of spins, in which the stick had been held forward in the first and then back, are reproduced in Figure 21. They are not greatly different for the two cases, but clearly show that, converted to full-scale equivalents, at 15,000ft any recovery of this aircraft had become impossible with an adverse moment of more than 10 to 15 units, whereas at 4,000ft it could be recovered in a time of less than 10 seconds for all moments applied, up to the highest value used of 30 units. Thus it was indicated that for the steep spin experienced with this type, the margin of safety (the additional adverse yawing moment which would cause recovery to become impossible) became significantly lower as altitude was increased.

Figure 21 Effect of equivalent altitude on spin recovery for biplane model (Reference 28)

209

Journal of Aeronautical History Paper No. 2015/03

Although the results were too few to lead to any simple generalisation, it is of greater note that these tests opened the way for further use of the method of applying adverse yawing moments in spinning trials with models. The origin of this procedure is somewhat obscured by the original idea having been first introduced in an internal RAE report by Gates, which for some reason was not selected by the ARC for publication in the R&M series (29). But a method based on this would soon take a place as an established feature in future routine testing in the vertical tunnel.

6.4 A potential set-back

Now that all necessary facilities seemed to be in place for making the comparisons of model and full-scale spins, it was feared that this part of the programme would be placed at risk through a change in the requirements of the Air Ministry (30). It had been decided that prolonged spinning would no longer be regarded as a regular Service manoeuvre. Accordingly, it was intended that the usual procedure at A&AEE, of allowing eight turns to take place, to ensure that the spin was steady before the controls were set for recovery, would be replaced by one requiring only two turns. This was considered to be more representative of events in service, where a trained pilot would be likely to attempt recovery as soon as it was realised that a spin had begun. The former procedure was to be retained only for training aircraft and certain small civil types.

At RAE a study was made of the possibility of reproducing the proposed new test procedure in a vertical tunnel (31). It would be necessary to include the initial, unsteady phases of the spin, as it could no longer be assumed that a steady state had been reached before recovery action began. Models were launched into the Free Spinning Tunnel from a trapeze in a straight stall or sharply-turning flight with wing-dropping to represent the manner of inducing the spin during full-scale testing. It was found that this initial motion could not be completed within the dimensions of the existing tunnel with models of a span of more than 1ft. To include these initial stages in routine tests with models of the size already found by experience to be needed would require a new vertical tunnel of about 30ft diameter. Moreover, analysis of the unsteady incipient spin had not been addressed sufficiently to give more than an outline of events in that phase. With the additional acceleration terms that would be involved there would inevitably be new uncertainties about how to relate model results to events at full scale.

After representations had been made by the ARC Spinning Panel, the Air Ministry agreed that although the revised procedure would be implemented as planned, additional spins as in the original test procedure would be included in trials as far as possible, in the interests of research. The Panel also pointed out that as model tests would continue to be made in the steady state, they could give no advance indication that there might be dangerous behaviour of an aircraft when recovery was attempted at an early stage. It was suggested that some device ought to be provided for the shorter test at full scale, by which it could be stopped promptly and safely if the aircraft was felt to be going out of control.

This suggestion was followed up at RAE in model tests made in the Free Spinning Tunnel and reported in early 1935 (32). Four possible devices were investigated – a parachute

210

Journal of Aeronautical History Paper No. 2015/03 attached near the tail, a spoiler on the outer wing tip, large fillets running along the top of the after part of the fuselage and a keel lowered at the side of the fuselage. The tail parachute, tested in a rudimentary form as shown in Figure 22, was found to be the most promising. It was confirmed that the parachute, flying clear of the wakes of the wing and tailplane, would be able to exert the full aerodynamic moment from its drag, and so pull the fuselage towards the preferred alignment to begin the recovery. For full-scale trials, it was expected that a location could usually be found for the parachute to be stowed internally near the tail, with means to deploy it quickly if the aircraft showed unexpected behaviour early in the spin. It might with advantage become a regular measure, available for use at any stage when the usual control actions for spin recovery were found to be ineffective.

It is not known for certain when a spin parachute was first used, but another occasion leading to its introduction occurred late in

1935, during the company trials of the Miles M.7 Figure 22 'Proof of concept' model test for anti-spin parachute Nighthawk. This was a low- (Reference 32) wing cabin monoplane with side-by-side seating and dual controls, being proposed as a training aircraft by Phillips and Powis Aircraft Ltd of Woodley, Reading (later the Miles Aircraft Company) (33). After spin recovery had been shown successfully with centre of gravity positions up to the design rear- ward position, the Air Ministry regulations required another test to be made, with the centre of gravity moved further aft by an amount equal to 10 % of the centre of gravity range. In this condition, recovery was found to be ‘difficult’, but the fuel tanks had not been full, so the company pilot Wing Commander F W Stent made a further test, with the tanks freshly filled. After 11 turns, the nose rose and the spin became flat. Nothing Stent tried had any effect, but after 23 turns without recovery, he managed to bail out, and the aircraft continued to spin to the ground and was destroyed. Given top priority at the works, the Experimental Department built a second prototype from scratch in 10 days and nights. F G Miles, who was Chief Test Pilot as well as Chief Designer, then personally completed the trials, this time successfully. Flight magazine reported later that ‘F G Miles has evolved a scheme for stopping a flat spin with the release of a small parachute anchored to the tail of the machine’ (34), but the extent of its use in those trials was not given. Then in 1936, the prototype of the Vickers Venom fighter was provided with a tail parachute for spinning tests at A&AEE, made in connection with the specification F.5/34. This was a version of the Jockey monoplane previously mentioned, which had also been destroyed in a flat spin. Although the design had been considerably modified, there had been a decision not to proceed with the trials, so it seems that no tests

211

Journal of Aeronautical History Paper No. 2015/03 with the parachute were carried out on that occasion. This safety feature was however widely adopted thereafter.

As reported earlier, the Bristol Fighter had become the work-horse of spin investigations at full scale and the vehicle for the development of the RAE standard instrumentation pack for recording the motion of the aircraft and its control positions. Tests at full scale and with models in free flight in the Balloon Shed had shown that this aircraft had two main stable spin states, a slow steep one and a fast flat one, but that its orientation was subject to large oscillations, which had been assumed to be indications of an instability that could cause transition from the one state to the other. In the latest work, one of the last studies involving a biplane, tests were made with a 1/20th scale model in the Free Spinning Tunnel for further comparisons to be made with the behaviour at full scale (35).

Results were obtained that showed both types of spin as before, and occasions when there would be changes from one type to the other. But there was also a third type, at an intermed- iate incidence and little bank, together with further irregularities of the motion. Inconsistencies in the observed behaviour occurred from day to day, despite the most rigorous attempts to control the conditions. It was concluded that this aircraft was extremely sensitive to small variations in the yawing moment, and further observations in the rotating balance at NPL would be needed before another attempt could be made at reconciliation of the results with those at full scale.

When these were included, detailed analysis had shown that the model did not exactly replicate the full-scale behaviour, but to a fair approximation it would have done so if there had been additional pro-spin moments in both roll and yaw of about 5 units. This lent further support to the view that the technique of attaching wing-tip vanes to apply additional moments would have wider application in model testing in the vertical tunnel.

6.5 The margin of safety

There was now a general framework, within which any results that appeared to be inconsistent with accumulated experience would become a focus of attention. As indicated in the previous Section, reconciliation of unexpected results from testing, particularly as between the Free Spinning Tunnel and at full scale, was now about to lead to a fundamental change of practice in the former.

There had been much concern about a few cases in which there had been a serious occurrence with an aircraft which had previously been found to spin and recover satisfactorily. It was thought possible that even small modifications to the airframe and changes to the weight and position of components, of a kind that would be expected to occur with any type during its time in service, might have led to failure to recover. There could be a marked difference in the nature of the spin, sometimes a progression to the flat spin, with its characteristic increase in incidence and rate of rotation. The occurrence of nutation in the motion around the spin axis was suspected of being evidence of instability at lower angles of incidence that was barely

212

Journal of Aeronautical History Paper No. 2015/03 restrained. It was worrying also that sometimes the change to the flat spin had become manifest only when the controls were moved to the position for recovery.

It was concluded that in the conditions of its original tests, the particular aircraft, though found to be satisfying the requirements, must have been close to a transition to a mode of spin that was potentially catastrophic. In the hope of detecting that, procedures for model tests in the vertical tunnel were extended so as to cover a wider range of effective altitude and possible variations in mass and its distribution, such as might occur in service. However, it was recognised that actual changes that might occur in practice were unknown and could not all be covered.

Thinking about the situation had suggested adopting routinely the procedure mentioned in Section 6.3 above - to modify the conditions of the test so as to introduce a defined ‘margin of safety’. It would now become the standard practice to worsen the spinning characteristics of the model deliberately by a known amount.

The argument for employing this approach had been given by Gates (29). The effects of a given mass distribution on the spin characteristics and recovery at full scale were represented when the scaling factors for mass and moment of inertia outlined in Section 3.5 were applied in the manufacture of the model. These accounted fully for the requirements of dynamic similarity of the motions of the model and the aircraft. But in respect of the aerodynamic moments, it was suspected that there could be additional effects of scale, going beyond the basic requirements for dynamic modelling. These might arise from possible differences between the patterns of the 3-dimensional airflow around the aircraft and its model, which could not at the time be assessed. From theory and practical experience it was now well known that the spin of an aircraft was more sensitive to changes in the yawing moment than to similar changes in either pitching or rolling moment. Hence it was assumed that an adequate margin in the test conditions to allow for possible aerodynamic scale effects could be made by the imposition of a yawing moment on the model, acting in a pro-spin direction.

Considering the use of a small vane attached to the tip of the inner wing to provide a known yawing moment, Gates showed that however the vane was orientated, it would make a contribution to the rolling moment as well. It was decided that a fair compromise would be o o to mount the vane so that its chord line made an angle of 120 to 130 to that of the wing. With a suitable choice of aerofoil section for the vane, the additional rolling and yawing moments could be about equal, but as the existing rolling moment was by far the larger, the proportional change in that would be relatively small when that of the yawing moment would be significant.

As the incidence at which a given model would settle in the spin was unknown beforehand, it was desirable that the aerofoil section chosen for the vane would give a resultant force curve that was fairly insensitive to changes in incidence beyond the stall. Suitable characteristics were found by using a symmetrical section based on the upper surface of the NACA aerofoil N.12. A vane of 4½in x 1½in planform was made with this section and the lift and drag characteristics were first measured in a wind tunnel in the ordinary way. Then it was attached to the lower wing of a 1/10th scale Bristol Fighter model, on a stalk that placed it to operate

213

Journal of Aeronautical History Paper No. 2015/03 at a distance of 1.4 times the semi-span from the fuselage axis in conditions that were well clear of wakes from the airframe. Measurements were then made of the yawing moment in o the NPL rotating balance. These showed that when set at 130 to the wing chord, the applied o yawing moment remained roughly constant for incidences up to 70 at appropriate rates of turn.

As a test of the proposed method, experiments were made in the spinning tunnel with this and several other models, including one of a single-seater fighter which had been found to have very satisfactory spinning characteristics (7). It was found that there was considerable variation between these types in the moment that could be applied before reaching a level at which recovery became impossible, with some models being able to recover satisfactorily with applied moments of up to 30 units. This encouraged the view that a test based on the method would discriminate well between those having good and bad spinning behaviour. That would be judged from the time taken for recovery when a standard additional adverse moment was applied.

It was decided to adopt a value of 10 units for further trials, about equivalent to that applied in normal flight by a typical rudder at full deflection. For a test to be considered satisfactory, the model would then be required to exhibit a stable spin, both with and without the vane in place, and to recover from that in both cases within an interval equivalent to 10 seconds at full scale following the movement of the control surfaces to the recovery position. The standard values of yawing moment and recovery time would be reviewed in the light of experience.

7 Effect of design features 7.1 Fuselage

The revised edition of AP 970 of May 1935, with Volume 1 now having the familiar title of Design Requirements for Service Aircraft, continued to include references to spinning, though only in connection with the tests that would have to be passed at A&AEE Martlesham Heath (36). As yet, there was no direction to sources of advice on design features that would assist in meeting the requirements for those, though more comprehensive results from the coordinated spin investigations were becoming available at this time.

In R&M 1535 by Irving reviewed above (18), early results were quoted from a lengthy study of the effects of fuselage cross-sectional shape on the yawing moment in the spin, obtained with models in the latest version of the rotating balance. The aim of this programme had been to investigate ways by which the anti-spin moments due to body and tail could be increased ‘without the adoption of any extraordinary features of design’. It was now completed and published as R&M 1689 (37). As the model was driven round by a motor in the rotating balance, conditions representative of fast and flat spins could be obtained, where they might not have been if only autorotative spinning had been available.

Wide differences were found in the yawing moment in the spin given by bodies with variations in profile shape and cross-section. Those with a profile that tapered towards a point at the rear were generally found to be inferior to others that retained some depth, as when

214

Journal of Aeronautical History Paper No. 2015/03 coming to a vertical stern post. Circular cross-sections were ‘peculiarly inefficient’, even to the extent of the rear fuselage tending to give a pro-spin moment at high yaw angles. Tests with a group of models having the same profile (one that retained some depth at the stern) showed clear variations with changes to the cross-sectional shape. For basically rectangular sections, the worst was with a semicircular part on the top of the rectangle, as in the ‘shellback’ arrangement that had been quite commonly used. The best was one with a semicircular part at the bottom, which in previous designs had often been flat. Plain rectangular and elliptical cross-sections were intermediate between those, the ellipse with the longer axis vertical being the better of these two. A keel along the top or bottom of the body had only a small effect, but strakes running along the top edges of the rear fuselage, equivalent only to about 1½ inch in width at full scale, gave big improvements with all sections tested. The Fairey Aviation Company was credited with introducing this feature, and it had been validated in spinning tunnel tests. In later years ‘anti-spin strakes’ of this nature were to become practically a standard modification tried out when poor spin behaviour was encountered.

It was possible to visualise the origins of many of these characteristics in terms of basic aerodynamics. For example, tapering the fuselage to a stern post rather than towards a point would add to the presented side area at the greater distances from the centre of gravity. This would increase the yawing moment obtained when a lateral component of flow was experien- ced due to sideslip in the spin. With a given profile, the effects of changes to the cross- sectional shape would be felt through variation in the side forces due to different drag coefficients when exposed to this lateral flow. Interruption of the flow around the top edge of the fuselage by strakes would also be expected to raise the drag coefficient for lateral flow.

Further tests were done to measure the moments with various arrangements of the empennage, an area of interest of long standing. For a given fin and rudder, blanketing them by the wake from the tailplane could render them ineffective in providing a restoring moment. The term now used for this was 'shielding' by the tailplane. It was found to be worst when that was mounted low on the fuselage, without much variation with changes to its fore-and-aft position because of the spread of its wake, as shown in Figure10. When the tailplane was mounted on the fin, shielding became progressively less as it was moved upwards, with the T-tail configure- ation being the best position, as had been noted previously. For a given area, the shape of tailplane made little difference, though changes to the aspect ratio over a range from 2 to 4 showed those of a higher aspect ratio to be slightly better, the wake at high incidence being somewhat reduced in width when the chord was narrower. For a tailplane in a typical location on the top of the fuselage, an important gain could be obtained by a change to its longitudinal position relative to that of the fin. This was especially marked with the fin mounted ahead of the tailplane, where it was not shielded at all by the latter’s wake. Most measured results of this kind could reasonably be interpreted within the current state of knowledge, though it had not always been easy to predict their relative merits prior to these tests.

Since measurements had been made either with the fuselage only, or with the fuselage and various arrangements of the empennage, it was decided to see the extent of changes that might occur if wings were present also. Two body shapes were used, those of the Bristol Fighter and of the Hawker Hornbill models, and results obtained with four wing arrangements; no

215

Journal of Aeronautical History Paper No. 2015/03 wings, biplane, low wing only and high wing only. The sequence of wing position given there is the descending order of yawing moment found in tests with the Hornbill fuselage, falling from - 40 to - 33.5 units full-scale. Although of sufficient magnitude to be taken into account, these differences were not thought serious enough to invalidate the conclusions of earlier tests made with different tail units. It was considered that the effect of wings was largely felt on the tailplane, and that with the various configurations the differences in the position of the wing wakes relative to that could be significant. A further effect was attributed to changes in the downwash there, as this could narrow the wake from the tailplane, resulting in some reduction in the shielding of the fin and rudder. Nevertheless, the investigators did not feel that their results would justify the expectations of the large effects of wing wake given by Korvin-Kroukovsky, reported section 4.2 (16). There had still not yet been any tests to map the wakes of wings and tailplane in the spin.

The tests with the Bristol Fighter model showed changes with wing interference that were generally in the same direction as those for the Hornbill, though somewhat greater in magnitude. Of particular interest, having regard to the spinning tunnel test results outlined in the previous o o section, is the report that ‘At about 50 and 60 incidences and at certain speeds of rotation very uncertain results were obtained, and there appeared to be a break in the continuity of the curves’. Thus it seems that the latest arrangement of the rotating balance was able to reproduce the complexity of the spinning behaviour of this machine that had been found with tests elsewhere.

7.2 Wings

Another report resulted from a continuing programme of work with models in the RAE Free Spinning Tunnel, now exclusively concerned with the low-wing monoplane configuration (38). For these tests, a standard fuselage and tail unit was used, as shown in Figure 23, fitted with wings of both rectangular and tapered plan-forms, but with the same area and aspect ratio. Details such as tip slots and split flaps were represented. The mass of the model and its distribution were changed so that systematic comparative tests could be made, with various combinations of the centre of gravity position and the dimensionless moment of inertia terms that were shown in sections 5.1.2 to 5.1.4 to determine the dynamic behaviour about the roll and pitch axes. Each of these terms contains the difference between the moments of inertia about two axes, respectively (B – C) and (C – A) in the form that appeared in the inertial moments. The justification here was that the spin state would be defined by these quantities when taken together with the yawing moment required to balance the third inertia moment. That was to be determined in the test programme for each condition by fitting wing-tip vanes, in the ‘margin of safety’ approach mentioned above. Vanes of increasing size were applied until the least size sufficient to prevent recovery against full opposite controls was found.

Typical results are shown in Figure 24, giving the time to recover from the spin versus the extra applied yawing moment for various conditions. In all cases the curves become increasingly steep as they approach the point at which recovery is no longer possible. Though the curves were inverted, this limit was termed the ‘precipice’, suggested by its abruptness. This indicated that having characteristics that placed a given design near to the

216

Journal of Aeronautical History Paper No. 2015/03

precipice would explain why some types of aircraft could be cleared in acceptance tests but found to be prone to spinning troubles when small changes had been made to them in service.

It was concluded from the overall results that the most important factor for low-wing monoplanes was the coefficient containing the difference (C – A) between the moments of inertia of the aircraft about the yaw and roll axes respectively. A large value of this would inhibit recovery, as can be seen in the upper part of Figure 24. It was determined mainly by the amount and distribution of mass in the fuselage. In a practical case of a heavily-loaded single-engined design, that term was likely to be large, especially where a long in-line engine was fitted. In that case, it was indicated that the fuselage and tail design would have to be very favourable to the production of a yawing moment if dangerous spinning was to be avoided.

On the other hand, small values of (C – A) seemed to be ‘very advantageous’, particular- ly with a forward centre of gravity position and a low value of (B – C), such as would be Figure 23 Model for effects of wing form on found if the loading was concentrated in the spinning of low-wing monoplanes wings. Thus, a twin-engine low-wing (Reference 38) monoplane would not be expected to spin dangerously even if the fuselage and tail designs were very unfavourable. These conclusions, now arising from tests with freely-spinning models, were consistent with the conclusions drawn from theory by Gates.

The effect of wing taper was found to be very small in all cases, but that of opening wing-tip slots was more complicated, varying with the mass distribution. It was unfavourable when (C – A) was large, but when this was small, opening the slot could change from slightly favourable to unfavourable as the centre of gravity position was moved aft. This ambiguous result was consistent with the variations found in earlier results from similar studies on the effects of slots in biplane wings, though there it was thought likely that all biplanes with tip o (39) slots would become unstable in roll if the incidence exceeded 45 . The overall position was summarised in the last report issued by the ARC Spinning Panel, after which spinning

217

Journal of Aeronautical History Paper No. 2015/03 was to be considered a part of the subject of Stability and Control and reported by the Sub- Committee of that name. It was stated briefly that ‘wing-tip slots, although they greatly decrease the liability to accidental spin, may increase the severity of a prolonged spin’ (40).

Figure 24 Effects of design features on spin recovery of models (Reference 38)

218

Journal of Aeronautical History Paper No. 2015/03

7.3 Spinning in the approach to WW2 7.3.1 Trends in design

The ‘monoplane revolution’ of the second half of the 1930s was slower to develop in Britain than in some other countries, but was given impetus by intelligence received about aircraft clandestinely developed for the formation of the Luftwaffe, which would soon be openly displayed. In due course, new types proposed for use in all branches of Service flying would be of this configuration, though not without false starts.

Air Ministry specification F.7/30 was intended to define the requirements for ‘zone fighters’ in the new role of Home Defence that was coming to the fore in strategic planning (25). This aroused the interest of manufacturers, but the requirements for an extended patrol duration by day and night, combined with a high enough maximum speed for engaging new designs of bombers, proved difficult to reconcile. There were also great problems with the Rolls-Royce Goshawk engine and its evaporative cooling (sometimes described as ‘steam-cooled’), which had been adopted by several of the contenders. As a result of many delays, competitive trials could not be held until mid-1935. And then no entry quite matched up to the requirements, though afterwards the Gloster entry was selected for production, becoming the Gladiator, the last biplane fighter to see service with the RAF.

The fitful progress of the F.7/30 competition caused urgent discussions to range over exactly what characteristics would be required for the next generation of RAF fighters. When a further specification, F.5/34, was issued, it concentrated on the daytime interception role, but in most respects was technically even more challenging. Requiring great steps to be taken in every area, this came near to visualising the archetypical interceptor fighter of the beginning of WW2 – the highly-loaded single-seat monoplane of all-metal stressed-skin construction, with high rate of climb and maximum speed, flaps and retractable undercarriage and an enclosed cockpit with oxygen supply. The fire-power of eight Browning 0.303in machine guns was reckoned to be needed to obtain a sufficient hit-rate to disable targets in the very brief periods of engagement expected.

Some interesting entries were designed for this specification, though in the event none of those submitted was considered to meet its perceived needs in full. One of them was the Bristol Type 133, to which reference was made earlier (see Figure 20). Another that might be noted here was the design by Henry Folland for Gloster Aircraft, shown in Figure 25a Aircraft with fin ahead of tailplane: Gloster F.5/34 (FAST Museum collection)

219

Journal of Aeronautical History Paper No. 2015/03

Figure 25a. This was to use the 840 HP Bristol Mercury IX nine-cylinder radial engine, which being air-cooled had not suffered the troubles of the Goshawk. The prototype aircraft K.8089 was not named, being known only by its company number and that of its specification F.5/34. Its monoplane configuration and novel features showed that even firms with a long history of supplying biplanes were adapting to the new circumstances. With others, this aircraft was transferred to RAE after the competition ‘to obtain research information’, an arrangement generally welcomed by the companies for the detailed test results that would be fed back to them subsequently. This case will serve to show that close support was already being given to designers on the types offered by the industry at this critical time, despite their inability to match the requirements of the latest specifications in full.

Extensive model spinning trials were made on the Gloster F.5/34 in early 1936, but the recovery characteristics of the design as originally submitted were deemed to be unsatisfactory (41). As with the Vickers Jockey mentioned above, the trials were then continued, with substantial modifications being made to the model, including 11 different arrangements of the empennage being tried. The most arduous of the test cases for the model had represented an equivalent operating height of 14,000ft, with the fuselage loaded to increase the moment of inertia B in pitch by 20 % and the centre of gravity placed at 6 % mean chord aft of the normal full load position. Then with 10 units of applied pro-spin moment the equivalent recovery time with the recommended tail modification was found to be 6.3 seconds, or 8.2 seconds with flaps and undercarriage lowered.

Gloster was heavily involved in production of the Gladiator, so that further work on the F.5/34 proceeded only slowly. The form seen in Figure 25a is that of the second prototype, incorporating the recommendations finally made by RAE, which included a lengthened rear fuselage, with the fin mounted ahead of the tailplane. In due course, the aircraft came back to RAE in 1939 for handling tests, when it received a good report, concluding with the statement that ‘The aeroplane is in general very pleasant to handle and is free from tricks’ (42). But by that time, the RAF was committed to other types for the interceptor role.

The distinctive arrangement with the fin ahead of the tailplane had been employed specifically to reduce shielding of the fin and rudder in the spin. This was the most overt design feature to arise from spinning research and communication of the results to the industry. It appeared in other aircraft at this time, notably the 2-seat carrier-borne dive-bomber produced under O.27/34 by Blackburn Aircraft, shown in Figure 25b, which entered service in October 1938 as the Skua, and in subsequent designs from the same company. It was also used again at Gloster, including in George Carter’s design for the E.28/39, Britain’s first jet-propelled aircraft (43), and for the DH 98 Mosquito, which was conceived in 1938/9.

220

Journal of Aeronautical History Paper No. 2015/03

Figure 25b Aircraft with fin ahead of tailplane: Blackburn Skua (FAST Museum collection)

7.3.2 The Hurricane and Spitfire

Hawker and Supermarine were already working on private ventures in the direction specified in F.5/34, and although they did not submit entries, they continued design work independently, with Air Ministry encouragement. Sydney Camm’s Hurricane and Reginald (‘R J’) Mitchell’s Spitfire which emerged from this were later to acquire iconic status, arising from their vital role with the RAF in the Battle of Britain in 1940. As their designs matured, prototypes were ordered with individual specifications based on similar requirements to those of F.5/34, becoming F.36/34 for the Hurricane and F.37/34 for the Spitfire. There had been close cooperation with Rolls-Royce during their development, and these aircraft would now be powered by the in-line V-12 liquid-cooled engine that was to become the famous ‘Merlin’. The prototype machines first flew in November 1935 and March 1936 respectively. It was a mark of the urgency with which high-performance fighters were required that production orders for 600 Hurricanes and 310 Spitfires were placed on the same day, 3rd June, 1936. In these orders, full armament and other improvements, embodied in a notional specification F.10/35, were to be introduced. Though the contractor and service trials which followed would inevitably focus on their performance and stability as platforms for their concentrated armament, their spinning characteristics soon came to be investigated.

The first model spinning tests of the ‘Hawker Monoplane’ had in fact been made just before the first flight of the prototype (44). In the same worst case conditions as those reported above for the Gloster F.5/34, it was reckoned to be on the verge of non-recovery, though there was little further effect when the flaps and undercarriage were lowered. It was thus considered to

221

Journal of Aeronautical History Paper No. 2015/03 be a borderline case, and it was recommended that the first full scale spinning trials should be made with a wooden propeller in place. This would keep the moment of inertia difference (C – A) as low as possible (it had been hoped to fit a metal two-pitch propeller to production aircraft, that would be appreciably heavier).

A long spin recovery time was confirmed in contractor’s trials of the Hurricane, so further model tests were authorised, ‘to determine the minimum modification which would make the spinning characteristics satisfactory’. These took place at RAE in December 1936 and (45) November 1937 . Attention was again directed to the tail of the aircraft, and a number of potential solutions evolved. The modification finally recommended (and adopted) was to fit a supplementary fin 6in deep along the bottom of the rear fuselage, extending forward beyond the tailwheel position, and to enlarge the rudder downwards to blend with that. By the time of the later tests, the standard worst case had been strengthened recently, to increase the applied pro-spin moment to 15 units. With the recommended modifications, the recovery time was then 15 seconds, though with 17 units recovery became impossible. This situation was generously rated ‘borderline’ but it was ‘anticipated that the recovery on the full scale aeroplane may also be slow’. Hurricane L1547 arrived at A&AEE for acceptance trials in June 1938, having been fitted with an anti-spin parachute as a precaution. However, spin recovery was reckoned to be normal, except that if the rudder was held on too long, there was a tendency for the aircraft to flick into another spin in the opposite direction (46).

The corresponding model spinning trials for the Spitfire were also begun before the prototype aircraft was flown, and its characteristics were again considered to be unsatisfactory (47). Experiments with modifications to the model resulted in a recommendation that the rear fuselage should be lengthened by 9in and the tailplane raised 7in, with which recovery became just possible at 10,000ft equivalent altitude in the worst test case, at a time when the standard pro-spin applied yawing moment was still 10 units.

Jeffrey Quill, who did much of the contractor’s test flying on the prototype, reported that the first spinning trials were approached with some trepidation (48). A crude anti-spin parachute arrangement had been fitted, stowed in a plywood box inside the cockpit. A cable ran from that along the top of the rear fuselage, held down to the skin only by adhesive tape, to a load- carrying fitting near the fin. In an emergency the pilot was required to seize the canopy and throw it over the side of the fuselage, where as it began to inflate it would rip the cable away from the fuselage until it deployed fully to provide the required moment for recovery. Having done its job, it could be released from the fitting by pulling a trigger in the cockpit.

With the parachute in readiness, Quill performed a full sequence of spinning trials with the prototype, covering the entire range of centre of gravity position, with spins in both directions and completion of eight turns in each before moving the controls for recovery. The aircraft was recovered without difficulty in every case. However, he reported that ‘K5054 always seemed to make a great fuss about it all while the spin was actually in progress’. This ‘fuss’ consisted of violent pitching oscillations of large amplitude, accompanied by variation of the rate of rotation within the turn, giving the spin the nature of ‘a series of convulsive flicks’. When something resembling this kind of nutation had been seen occasionally in the early modelling research (though never of such an extreme extent), it had been taken to be a

222

Journal of Aeronautical History Paper No. 2015/03 dangerously unstable motion that would lead the aircraft towards taking up a fast flat spin from which recovery would be impossible.

Quill’s report did not prevent the transfer of the aircraft, with a proper anti-spin parachute installation, to A&AEE in late 1938 for acceptance trials, where the aircraft had been found to behave satisfactorily (46). These trials were the definitive ones, notwithstanding any reservations arising from the contractor’s testing. It was by then accepted that in full scale tests, this aircraft had not shown its spin to have tendencies that could be termed 'dangerous'. But model tests were resumed at RAE in the spring of 1939 (49). The objective now was research, by which it was hoped to ‘correlate model and full scale spinning characteristics and to determine the minimum modification which would make the model satisfy the revised spinning standard’ (i.e. with the requirement of an applied moment of 15 units). Many variations of the rear fuselage and empennage were tried, with the conclusion that for the best result in this situation the fuselage should be lengthened by 12in and the tailplane raised by 10in. Then under the worst conditions of loading, it was predicted that recovery at an equivalent altitude of 15,000ft could be achieved in 10 seconds full scale against 13 units of yawing moment. This was still a borderline case, since recovery was impossible if the full 15 units were applied.

It was evidently thought that correlation between model test results and full scale testing was not yet nearly good enough to justify changes to the design at this critical time, and the Spitfire was not modified in the way recommended. It subsequently fell to Quill to carry out the initial contractor’s spinning trials on every mark of Spitfire and of its successor the Seafire. He reported that over the next decade, there was ‘never a case of any production mark of either failing to recover from a spin’ (48).

7.3.3 An urgent review

The adoption of the raised standard for the applied yawing moment and the minimum recovery time in model spinning tests had been the outcome of a review by Gates (50). To make a general comparison between model and full-scale tests, he assembled evidence for the latter from reports of contractors’ trials, official A&AEE trials and special tests at RAE. For these cases the mass and moments of inertia would be known, but in the corresponding model tests these were to be enhanced to take account of changes that might credibly occur up to the point of entry into service and perhaps beyond. It was the usual aim to do model tests before the first flight of the prototype, and there was then a much greater range of possible future variations and combinations, so that it was prudent for wider margins to be applied.

There were often cases where model tests before first flight had indicated the need for modifications to the design to obtain satisfactory spin characteristics. Suggestions had then been made for minimum changes that had at least brought the case into the borderline area, and where these were implemented the tests at full scale were usually acceptable. As in the case of the Spitfire, the history of the type in service had been sometimes so much better as to indicate that a larger margin of safety had been obtained in practice than that indicated by the model tests. There were however, some more worrying examples, such as that of the Bristol Type 133, mentioned earlier, for which pre-flight model tests and the contractor's pre-flight

223

Journal of Aeronautical History Paper No. 2015/03 trials had been satisfactory, but when a spin was made with the undercarriage lowered the aircraft had entered a flat spin and had been wrecked.

Another case causing concern was that of the Philips and Powis Hawk Trainer, which from 1936 was being considered for the role of the RAF's first monoplane ab initio training aircraft (33). This and other types in the Hawk series had not experienced any problems with recovery from spins, and no pre-flight model tests had been done. But at the end of its contractor’s trials there had been a failure to recover with maximum weight and full aft centre of gravity position and this occurred again when one went for tests at Martlesham Heath, that time fatally for the RAF pilot. A further test, which would otherwise have also been irrecoverable, was ended safely as the aircraft had fortunately been fitted with an anti-spin parachute. This was another of the earliest reported uses of this safety device.

Subsequent model tests at RAE gave a very satisfactory recovery curve, without reproducing the behaviour that had led to the crash. But some now familiar advice was given that the tailplane should be raised by 6in and the top of the rear fuselage flattened and fitted with strakes ahead of it. These modifications, together with the fitting of a taller rudder for reasons not connected with spinning, were made to subsequent production aircraft and retrospectively to earlier ones, and further spinning tests at full scale were carried out without trouble. In extensive service as the Magister, its uses included pilot training in spinning and recovery, and there were some incidents in which spinning was implicated, but no further action had been required.

Unexpected spinning incidents, especially of a nature that had not been reproduced in model testing, were troubling. Gates pointed out that no practicable test could be guaranteed to cover a situation that might occur just once in the lifetime of the aircraft. The number of possible combinations of changes of mass and its distribution and of the positions of controls, flaps, slots and undercarriage was very great and much reliance must be placed on the experience of the experimenter in focussing on those that were thought to be the most likely to be trouble- some. The possibility would always remain of overlooking the significance of some combin- ation of features of any given design, perhaps with atmospheric conditions which could not be reproduced in the tunnel. It would be hoped that the accumulation of test results over time would give indications that suggested where previously unexpected behaviour might be likely to arise. Meanwhile, he proposed that the pass criteria of the model spinning test should be strengthened, so that the applied adverse yawing moment would be raised from 10 to 15 units. On the other hand, he considered that the time allowed to recovery, after the controls were moved to the usual positions, could be increased to 10 seconds full scale from the previous value of 8 seconds. These changes were duly adopted, while other practices, on the range of the centre of gravity position and increases in moments of inertia, continued unchanged.

A need for an increase in tunnel speed had now become apparent, and this was put in hand. With monoplanes had come wing loadings that were much higher than those around which the Free Spinning Tunnel had been designed, and these were expected to rise further. The basic scaling laws in section 3.5 show in equation 1 that the speed of descent of an aircraft in a spin increases with the square root of the wing loading. Then from equation 2, if the tunnel

224

Journal of Aeronautical History Paper No. 2015/03 speed is fixed, correct modelling requires that for a higher wing loading the corresponding model must be made to a smaller scale (larger n). But there were practical limitations to the smallest size of models if the necessary features of the design were to be adequately represented. These now included flaps, slats and retractable undercarriages, and it was thought that finer detailing of control surfaces was needed if behaviour in the recovery phase was to be correctly reproduced. Of necessity, some models had been made with a span of only 12in, but the over-arching concern about the unknown ‘scale effect’ argued strongly against anything smaller than about twice that size.

It had been necessary also to make compromises with regard to the requirement for the equality of the ‘weight density ratio’ on which the dynamic scaling of tunnel models was based (see section 3.5). If the desired value of this could not be obtained, it would not necessarily mean that the result of the tunnel test was totally invalid. Usually, the result would correspond to the full-scale behaviour at a lower altitude than was customary for the aircraft’s spinning trials. Some examples of where this had occurred have been cited above.

The need for compromises such as these was reduced by increasing the maximum air speed in the tunnel. With a new drive motor of 120HP fitted in 1938, this was raised from 35 to 56ft/s.

8 A significant staging-point 8.1 Basic principles

It was seen that in the closing years of the 1930s the ‘margin of safety’ approach to model spinning trials at RAE included the provision of advice to designers on improvements to the spinning characteristics as required, based on tests made with modifications to the model. But efforts continued with a new variant of the approach that had aimed in the past to give designers the tools to ensure that satisfactory spinning behaviour could be ‘designed in’ as a new type evolved. The objective had long been to identify the features of an aircraft that would lead to good spinning behaviour and to obtain specific values or expressions to represent them, with which the designer could routinely calculate the dimensions required. This had not been achieved so far, but it was hoped that progress could now be made through an analysis, repeatedly up-dated, of the accumulating results of spinning tests of all kinds. The best outcome would be that over time a set of design rules would at last emerge in a suitable form for use at all stages in the design process.

In essence, the outcome of a spin test was one of ‘pass’ or ‘fail’. The matter now was to see if these bald results could be analysed in a systematic way that would not simply indicate which general features of the designs were favourable or not, but would allow numerical data to be extracted about the location of the boundary between the two possible outcomes. Though that might seem to be a largely empirical procedure, it was firmly underpinned by theory, which indeed had informed progress in the subject from the beginning.

It was understood from early times that the spin was an interaction between two basic components, the inertial and the aerodynamic. In later work it had been concluded that the

225

Journal of Aeronautical History Paper No. 2015/03 most important elements in these were respectively the difference of moments of inertia (C – A) and the yawing moment that could be applied by the rear fuselage and empennage in sideslip when in the spin. These were the main determinants of the angles of incidence and yaw taken up by the aircraft relative to its helical path, and of the rate of rotation about the vertical axis of the path. In the light of these considerations, it might be conjectured that the values of two factors, in which these elements were represented, could help to define boundaries between designs with a propensity to pass or fail the spinning tunnel test. Perhaps just these two factors would suffice to determine the basic character of the spin. But for any model that had passed, the rudder must have been able to apply an anti-spin moment sufficient to slow the rotation, the necessary first step in bringing the aircraft out. Thus, the possibility of recovery would depend on a third factor, probably taking some form that resembled what had become known as the rudder volume in the treatment of lateral stability.

8.2 Definition of the terms

Parts of the proposition summarised above had been under discussion for some time, but the (51) definition of preferred terms was now addressed in R&M 1810 of July 1937 . It was acknowledged that the general layout of an aircraft is largely determined during early stages of design by the terms of the operational requirement. As significant changes to the moments of inertia can be made only by moving components of appreciable mass from one part of the airframe to another, there are fewer opportunities for making these as a design becomes consolidated. And so, as soon as the (C – A) term could be reasonably estimated, it could provide a kind of base-line quantity, representing the essential inertial features of the particular aircraft. Then the later decisions in design concerning its behaviour in the spin would be to ensure that the side areas of the fuselage and fin were sufficient to generate a yawing moment that could restrain the motion from developing into a flat spin. And finally, the position and area of the rudder must enable it to begin the recovery when the pilot actuates it.

To define these controlling elements in numerical terms, they can be made dimensionless by dividing them by appropriate groups of the leading dimensions of the aircraft. If these groups are chosen on the same basis as those for determining the scale factors for model spinning outlined in section 3.5, the dimensionless forms will have the same magnitudes for an aircraft and for a model of it. During the initial work on this at RAE, the three resulting terms and their definitions had emerged as follows (NB no standard symbols for the terms had been given and X, Y and Z are assigned here for the current use only) :

3 Inertia coefficient X = (C – A)/ ρ Ss (8)

Body damping ratio Y = Σ (dA x 2)/Ss 2 (9) f

Unshielded rudder volume coefficient Z = A x /Ss (10) r r

In equation 9, dA is the presented area of an element of the fuselage and fin in side view, f and x is the distance from the centroid of that element to the centre of gravity. In equation 10,

226

Journal of Aeronautical History Paper No. 2015/03

A is the presented side area of the rudder and x is the distance of its centroid from the r r centre of gravity. (These two terms are similar to others arising in the calculation of lateral stability, and their formations would be familiar to designers. Equation 9 had already appeared in a similar form in equation 7 in Section 5.1.5. The description ‘volume coefficient’ in equation 10 arises from the numerator being the product of an area and a length, which has the dimension of a volume, though no physical volumetric element of the aircraft is actually involved). The air density ρ required in equation 8 is that for the full-scale test as represented in the scaling of the model.

The aim of the first study was to obtain minimum acceptable values of damping ratio and rudder volume by analysis of the test results available from the vertical tunnel. Aircraft types coming for test, at full scale and as models, and apparently having generous provision of side areas, had nevertheless been found to have unacceptable and sometimes catastrophic spinning characteristics. This had long been attributed to reductions in the ‘effectiveness’ of an element of side area if it had been shielded from the airflow in the spin by other parts of the airframe, especially when in the wake of the tailplane. A vital new part of the procedures now being developed was that the areas dA and A were to be factored according to the extent to which f r they were shielded in this way. It is evident that the methods used to estimate the factors would be critical to progress along this path. The criteria for assigning those were not immediately apparent, so it was expected that there would be lessons to be learnt about the most effective values from experience.

8.2 Representation of shielding

R&M 1810 seems to have been the first published account in which proposed shielding factors are represented (51). This gives an analysis of tests in the vertical tunnel, exclusively on models of monoplanes ‘in view of the trend of modern design’. Results for twenty-two types were available, obtained using the applied yawing moment technique. In these tests, the criteria for passing were those prior to the modification proposed by Gates (50). Only six designs had passed without modification, though two had been submitted with alternative tail arrangements, both of which proved satisfactory when those were represented. A further 14 designs were passed on re-testing after being modified, though two others were considered to be ‘dangerous’ and there were no further tests of those. Judgments were made for cases that were deemed to be ‘borderline’ as to whether they might reasonably be assigned to either ‘pass’ or ‘fail’.

The method of allowing for shielding by the wake of the tailplane for the calculation of the body damping ratio and the unshielded rudder volume is illustrated in Figure 26. As appropriate for the first use of the procedure, it was kept very simple. Areas above the o tailplane are assumed to be shielded by it if lying within an angle to the fuselage axis of 60 backwards from the leading edge and 30 o from the trailing edge. This is intended to represent o o the situation in a spin with an incidence of 45 and to make an allowance of 15 for divergence of the edges of the wake (this would be roughly in accordance with results such as those shown in Figure10, though the work from which those were derived is not cited).

227

Journal of Aeronautical History Paper No. 2015/03

Figure 26 Definition of Body Damping Ratio and Unshielded Rudder Volume for first use (Reference 51)

Areas that are shielded within the boundaries of the wake are then omitted from the calculations entirely. This is equivalent to deciding that areas within those limits produce no significant yawing moments, while areas outside them are fully effective. On the other hand, there had been some evidence that the side area of the fuselage beneath the tailplane became more effective by there being a pressure rise in that location, so a factor of two is applied to that area. Because of the great variety of empennage arrangements in the types tested, which included for example twin tails, the definitions of the areas considered to be shielded are adjusted as required, but along similar lines.

The results are first presented in two plots of the body damping ratio and the unshielded rudder volume coefficient thus obtained against the inertia coefficient as the independent variable. Over the range of the latter, the variation of the two moment coefficients for the different aircraft is surprisingly great, in the case of the rudder volume covering almost an order of magnitude. But in neither case does the plot show a clear distinction between those points representing the ‘pass’ and ‘fail’ cases, as had been hoped. The points for the two cases are seemingly scattered and intermingled indiscriminately. So in this first attempt, it proved not to be possible to place boundary lines on the plots which would separate them, and they are not being reproduced here.

228

Journal of Aeronautical History Paper No. 2015/03

It seems almost in desperation that another plot is tried, of the product of the moment coeffic- ients, as no justification for doing this is offered, and the form of that product does not appear to have any physical significance. The reason for including it must be just the finding that such a plot does allow a line to be drawn which separates the ‘pass’ results from the ‘fails’ fairly well, though the number of results is too few to allow any conclusions to be reached. The best advice that could be given from this exercise was that designers should ensure that the value of this product for the aircraft well exceeds the one obtained for the appropriate point along the dividing line on this plot, according to the value of its inertia coefficient. Preferably, it was advised, that the value should be twice as great.

This was not a satisfactory outcome, and could not be commended to designers. But evidently regarding it as ‘work in progress’ the ARC Stability and Control Sub-committee thought the approach to be promising, and following events were to confirm that.

8.3 Refinement

It becomes more difficult to locate details of work done at this time and subsequently. Publication of the more significant research findings in the form of R&Ms became caught up in preparations for war. ARC Technical Reports were issued for 1938 and for each year of WW2, though later they would be without the usual technical summaries prepared for the information of the Secretary of State for Air. There was now an additional administrative layer checking papers for security, and through this and other delays not all were ready when the relevant volume had to go to press. Four Special Volumes were published after the war, to contain the R&Ms that had been intended for publication but had missed the deadlines. Though this process seems rigorous enough, it is found that the volumes for 1938, 1939 and 1940, covering the remainder of the period considered here, contain few papers on spinning, and none appears in the post-war Special Volumes. Accordingly, the RAE Reports and Technical Notes are the main sources for that time. These have been reviewed and some are already cited above where appropriate, but it seems unlikely that systematic recording of work could have proceeded normally under the circumstances. (It has been noted elsewhere, for example, that model spinning tests were carried out in early 1940 for the development of the first British turbojet aircraft, the Gloster E.28/39, but the results had not been written up in an RAE report until 1944 (43))

In the last of the years covered by this paper, work continued on the approach outlined above, turning it eventually into what amounted to a new standard procedure for the assessment of the propensity of an aircraft to spin dangerously and the ability to recover it. With further elaborations in the light of experience, this general form would stand for the duration of the war and beyond. An up-dated report on the procedure of May 1940 from the Airworthiness Department at RAE provides an appropriate basis on which to conclude the present review (52).

The main advance here is a refinement of the simple weighting given initially to lateral areas in the calculation of the Damping Coefficient and the Unshielded Rudder Volume Coefficient (51). In making comparisons of numerical data from different times, it should be noted that the characteristic linear dimension used in the definitions of these terms is now the span b, instead

229

Journal of Aeronautical History Paper No. 2015/03 of the semi-span s used previously, as given in equations 8, 9 and 10. Not much is recorded about the method by which the new weighting factors had been determined, though that could have involved reference to existing data on yawing moments in spin situations from tests with the rotating balance, or made purely by ad hoc variation of the weightings until a set was found that produced the clearest distinction obtainable between the pass and fail results from the vertical tunnel tests. The area weighting factors (given the symbol n) now proposed are illustrated in Figure 27. Many of the original values are retained, including the enhanced contribution from the fuselage side area below the tailplane, though over part of that the factor n is reduced from 2 to 1.5. A new element is the area on the fin just above the tailplane, where the contribution to the Damping Coefficient has become negative, implying that there is suction at that point when in sideslip, with the weighting factor n taking a value of -0.25.

Figure 27 Side-area shielding factors (Reference 52)

The effects of these changes on the Damping Coefficient and the Unshielded Rudder Volume are shown in Figure 28 for the model results as previously used. Lines could now be drawn which separate distinct regions in which the great majority of the results are either pass or fail. Finally, in Figure 29 plots are shown for results of tests at full scale processed in the same way. As the number of available results is smaller, the method used to draw the boundary lines was to give them the same slopes as those found for the model tests, which are shown dotted in the figures.

230

Journal of Aeronautical History Paper No. 2015/03

Figure 28 Unshielded rudder volume and damping coefficients, model results (Reference 52)

231

Journal of Aeronautical History Paper No. 2015/03

Figure 29 Unshielded rudder volume and damping coefficients, full scale results (Reference 52)

232

Journal of Aeronautical History Paper No. 2015/03

It is seen that some quite convincing boundaries were then obtained for the full scale results, though in somewhat higher positions. This was encouraging for the continuing hope that results from tunnel and free flight tests would converge as more became available.

Simple linear equations are given for the boundary lines at full scale, providing the basis for straightforward routine checks for the use of designers. These take the following forms (the terms X, Y and Z are defined in equations 8, 9 and 10, section 8.1):

Damping Coefficient Y = 0.0125 + 0.106 X (11)

Unshielded Rudder Volume Z = 0.0070 + 0.024 X (12)

For a given value of the Inertia Coefficient X, which could be estimated fairly well even at an early stage in design, these equations give minimum values of Damping Coefficient and Unshielded Rudder Volume that should be adopted for satisfactory spinning behaviour and recovery.

In the plots of Figure 29, the positions of the data for the Hurricane and Spitfire might be noted; they are given by the reference numbers 16 and 18 respectively. These show that the values of their Inertia Coefficients X are very similar, around 0.15. Both Y and Z for the Hurricane lie just above the boundaries on the 'pass' side, while for the Spitfire Y is just below the boundary but Z (the Unshielded Rudder Volume) is definitely in the 'fail' area. However, no significant changes to the rudder were made for the various Marks of the Merlin- engined aircraft before the Mk XVI, so the low value of Z seems not to have caused any difficulties in recovery for this type (48).

The significance of this procedure is clear. Unlike the yawing moment coefficients used by Irving and others, which would have to be estimated from tabulated or graphical results from rotating balance tests, these coefficients involve only the physical dimensions of the aircraft and the area weighting factors n given by a diagram. Many more test results at full scale would be required to locate the boundaries more reliably. But even with the initial guidance now available, checks could be made as the design evolved on the adequacy of the side areas of the rear fuselage, fin and rudder to indicate the prospects of the aircraft having satisfactory spinning and recovery characteristics. Further, some assurance could now be gained before the start of company trials with the prototype or acceptance tests at A&AEE as to the safety of entering into the spin.

This method would form the basis of developments in spinning and spin research in future years, initially through further refinement of the calculation of the side area weighting factors. In due course further consideration was given to the possible effects of the ratio of the rolling and pitching moments of inertia A/B, as indicated from theoretical studies by Gates and Francis (19). This aspect came to have more prominence with the emergence of more aircraft types with wing-mounted engines, where the masses located laterally at some distance from the centre of gravity added significantly to A without a corresponding increase in B.

233

Journal of Aeronautical History Paper No. 2015/03

9 Conclusion

Throughout the period covered by the second part of the paper, research on spinning and spin recovery continued to be pursued at NPL and RAE by theory, model testing in rotating balances and flight testing at full scale. A further major step forward was made with the addition of the vertical Free Spinning Tunnel at RAE, which was used to test dynamic models of particular aircraft and also for systematic model testing to investigate the effects of design features on the nature of the spin and measures for recovery from it. With guidance from theory, a method of presenting data from model and full-scale testing was developed that defined boundaries beyond which three key dimensionless groups of aircraft properties should lie to provide good prospects of obtaining satisfactory spinning and recovery characteristics.

For the first time, this method provided designers with a straightforward procedure by which the requirements for the control of spinning characteristics could be included with other established routines in the development of a new design. The history of spinning and spin research in the UK was by no means at an end, but taking the long view, this point at the beginning of the 1940s could be seen as one of the more significant staging-posts on the journey.

______

Acknowledgements Much valued help with locating sources for this study was given by Brian Riddle and Chris Male, RAeS and Alan Brown and Peter Pearson, FAST

`

References 1 - (1931) Annual Technical Report of ARC for 1929/30 HMSO, London 2 Irving H B and Batson, A S (1930) Spinning experiments and calculations on a model of the Fairey IIIF seaplane with special reference to the effect of floats, tailplane modifications, differential and floating ailerons and “interceptors” ARC R&M 1356, HMSO, June 1930 3 Scott-Hall, S (1931) The spinning of aeroplanes J Royal Aeronautical Society, Vol.35, pp.608-623 4 Scott-Hall, S (1932) Aeroplane performance testing and practical aspects of methods employed at Martlesham with civil aircraft Aircraft Engineering, Vol.4, 1932, pp.112-114 5 Gates, S B (1931) Measured spins on Aeroplane H ARC R&M 1403, HMSO, April 1931

234

Journal of Aeronautical History Paper No. 2015/03

6 Gates, S B (1931) The effects of centrifugal force on the controls in a spin ARC R&M 1416, HMSO, May 1931 7 Irving H B, Batson A S and Stephens A V (1931) Spinning experiments on a single seater fighter with deepened body and raised tailplane. Part I Model experiments, Part II Full scale spinning tests ARC R&M 1421, HMSO, Dec 1931 8 Gates, S B, Ormerod, A, Fairthorne, R A, Stephens, A V, Irving, H B and Batson, A S (1932) Experiments on the Hawker Hornbill biplane ARC R&M 1422, HMSO, Aug 1932 9 Stephens, A V (1931) Free-flight spinning experiment with single-seater aircraft H and Bristol Fighter models ARC R&M 1404, HMSO, April 1931 10 Stephens, A V (1931) Free model spinning researches Aircraft Engineering, Vol.3, 1931, pp.213-215 11 Irving H B and Stephens, A V. Safety in Spinning J Royal Aeronautical Society, Vol.36, 1932, pp.145-204 12 - (1933) Annual Technical Report of ARC for 1931/32 HMSO, London. 13 Gates, S B (1931) The determination of the moments of inertia of aeroplanes ARC R&M 1415, HMSO, March 1931 14 Bryant, L W and Miss I M W Jones (1932) Notes on recovery from a spin ARC R&M 1426, HMSO, March 1932 15 Jones, B Melvill and Haslam, J A G (1932) Airflow about stalled and spinning aeroplanes shown by cinematographic records of the movements of wool-tufts ARC R&M 1494, HMSO, Aug 1932 16 Korvin-Kroukovsky, B V (1932) The uncontrolled tail spin Aircraft Engineering, Vol.5, 1932, pp.105-112 17 Petersohn, E (1932) Wake measurement behind wings in separated flow (in German) Zeit für Flugtechnik u Motorluftschiffahrt, Vol.22, Part 10, 28 May 1932, pp.289-300 18 Irving, H B (1933) A simplified presentation of the subject of spinning of aeroplanes ARC R&M 1535, HMSO, March 1933 19 Gates, S B and Francis, R H (1934) An analytical survey of the effect of mass distribution on spinning equilibrium ARC R&M 1644, HMSO, Sept 1934. 20 - (1935) Technical Report of ARC 1933/34 HMSO, London 21 Stephens, H V and Cohen, J (1932) Spinning of Pterodactyl Mark IV ARC R&M 1576, HMSO, July 1932 22 Martin, C (1988) The spinning of aircraft – a discussion of spin protection techniques including a chronological biography Dept. of Defence, Commonwealth of Australia, ARL Report Aero-R-177, Aug 1988 23 Allwork, P H (1933) A continuous rotation balance for the measurement of yawing and rolling moments in a completely represented spin ARC R&M 1579, HMSO, Nov 1933

235

Journal of Aeronautical History Paper No. 2015/03

24 Halliday, A S and Burge, C H (1934) Experiments on the Whirling Arm ARC R&M 1642, HMSO, August 1934 25 Sinnott, C (2001) The Royal Air Force and Aircraft Design, 1923 – 1939 Frank Cass, London 26 Irving, H B, Batson, A S and Gadd, A G (1933) A comparison of the spinning properties of high and low wing monoplanes obtained from autorotation experiments (with yaw) ARC R&M 1534, HMSO, Feb, 1933 27 Stephens, A V and Francis, R H (1933) Model spinning tests of an Interceptor Fighter ARC R&M 1578, HMSO, May 1933 28 Gates, S B and Stephens A V (1934) The air density effect in spinning ARC R&M 1663, HMSO, Oct 1934 29 Gates, S B (1934) A method for providing for scale effect in model spinning tests RAE Report BA 1154, Royal Aircraft Establishment, Sept 1934 30 - (1936) Annual Technical Report of ARC for 1934-35 HMSO, London, pp.33-36 31 Gates, S B and Francis, R H (1934) Representation of a normally entered spin in the free spinning tunnel RAE Report BA 1155, Royal Aircraft Establishment, Sept 1934 32 Francis, R H (1935) Safety devices for full scale spinning trials RAE Report BA 1195, Royal Aircraft Establishment, April 1935 33 Amos, P (2009) Miles Aircraft – the early years Air-Britain (Historians) Ltd, Tonbridge 34 - (1936) Spin-stopping by parachute Flight, Vol.50, 13 Feb 1936, p.187 35 Alston, R P and Cohen, J (1936) An analytical comparison of model and full scale spinning experiments on a Bristol Fighter ARC R&M 1726, HMSO, March 1936 36 - (1935) Design Requirements for Service Aircraft Air Publication 970, Air Ministry, HMSO, London 37 Irving, H B, Batson, A S and Warsap, J H (1935) The contribution of the body and tail of an aeroplane to the yawing moment in a spin ARC R&M 1689, HMSO, Nov 1935 38 Francis, R H (1936) Interim report on systematic model research in free spins: low wing monoplanes ARC R&M 1714, HMSO, Jan 1936 39 Irving, H B, Batson, A S and Warsap, J H (1935) Spinning experiments on a model of the Bristol Fighter aeroplane, including the effect of wing tip slots and interceptors ARC R&M 1654, HMSO, Feb 1935 40 - (1937) Annual Technical Report of the ARC for 1936 HMSO, London, pp.23-24 41 Finn, E, Alston, R P and Francis, R H (1936) Model spinning tests of Gloster F.5/34 RAE Report BA 1276, Royal Aircraft Establishment, Feb 1936

236

Journal of Aeronautical History Paper No. 2015/03

42 Morgan, M B (1939) Research handling tests on Gloster F.5/34, K.8089 RAE Report BA 1523, Royal Aircraft Establishment, Feb 1939 43 Brinkworth, B J (2012) The Gloster E.28/39 – Fin arrangement and spinning characteristics J Aero Hist, Vol.2, 2012, pp.65-82 44 Alston, R P and Shone, H G (1935) Model spinning tests of Hawker Monoplane (P.V. to Specification F.36/34) RAE Report BA 1229, Royal Aircraft Establishment, Sept 1935 45 - (1938) Further model spinning tests in the Hurricane RAE Report BA 1229a, Royal Aircraft Establishment, May 1938 46 Mason, T (1993) British Flight Testing, Martlesham Heath, 1920 – 1939 Putnam, London 47 Alston, R P and Shone, H G (1935) Model spinning tests of Supermarine F.37/34 RAE Report BA 1225, Royal Aircraft Establishment, Aug 1935 48 Quill, Jeffrey (1983) Spitfire John Murray (Publishers) Ltd, London 49 - (1939) Further model spinning tests on the Spitfire RAE Report BA 1532, Royal Aircraft Establishment, April 1939 50 Gates, S B (1937) Note on model spinning standards RAE Report BA 1436, Royal Aircraft Establishment, Oct, 1937 51 Finn, E (1937) Analysis of routine tests of monoplanes in the Royal Aircraft Establishment Free Spinning Tunnel ARC R&M 1810, HMSO, July 1937 52 Tye, W and Fagg, S V (1940) Spinning criteria for monoplanes RAE Report AD 3131, Royal Aircraft Establishment, May 1940.

237

Journal of Aeronautical History Paper No. 2015/03

Appendix 1 The ARC standard definitions and nomenclature of 1936

238

Journal of Aeronautical History Paper No. 2015/03

239

Journal of Aeronautical History Paper No. 2015/03

The author

Brian Brinkworth read Mechanical Engineering at Bristol University. He worked first on defence research at the Royal Aircraft Establishment Farnborough during the 1950s. There, he was assigned part-time to be Secretary of the Engineering Physics Sub-Committee of the Aeronautical Research Council (ARC), and after moving into Academia in 1960, he was appointed an Independent Member and later Chairman of several ARC Committees and served on the Council itself. Thereafter he was appointed to committees of the Aerospace Technology Board.

At Cardiff University he was Professor of Energy Studies, Head of Department and Dean of the Faculty of Engineering. For work on the evaluation of new energy sources he was awarded the James Watt Gold Medal of the Institution of Civil Engineers. In 1990 he was President of the Institute of Energy and elected Fellow of the Royal Academy of Engineering in 1993.

Since retiring, he has pursued an interest in the history of aviation, contributing papers to the journals of the RAeS, which he joined in 1959. He holds a Private Pilot’s Licence.

240