CONCORDE SST - HISTORY

EARLY DEVELOPMENT

Early Development

When, in November 1962, the British and French governments agreed to develop and build a supersonic airliner, they and the manufacturers knew that there would have to be an exhaustive research and development programme before the aircraft could be certificated for passenger operation. Even those closest to the project did not at that time foresee the full-scale complexity and cost of this programme.

The Concorde engineering team was working on the frontiers of technical knowledge and preparing to venture into areas not hitherto explored by commercial aircraft designers. They could not know where their research would lead them or what unpredictable problems might be ahead. In the event, the Concorde test programme was to require more than a decade of research on the ground and nearly 5,000 hours of flight development, by far the most thorough and comprehensive programme ever mounted in support of a civil aircraft type.

In the earlier years of the programme the research effort was concentrated principally in aerodynamics, materials and structures. While this was going on, the engineering organisation faced up to the difficult task of formulating a preliminary aircraft design, sufficiently detailed to enable marketing discussions with potential customers to be started.

Aerodynamics and materials

Every supersonic aircraft design presents the aerodynamicist with a range of difficult problems, including two which are of supreme importance. One is the aerodynamic aspects of the powerplant installation. Propulsive efficiency is a critical factor in subsonic aircraft design, but it even more crucial to the success of a supersonic aircraft and more difficult to achieve because of the widely varying airflow demands of the engine in different phases of a supersonic flight.

The second challenge arises, not from a specific installation, but the total airframe configuration. To meet it the aerodynamicist has to produce a satisfactory compromise between two inherently conflicting requirements: the need for minimum drag in supersonic flight and the need for controllability and ease of handling in subsonic flight, particularly in landing and take-off. Trim tabs, spoilers and other external moving surfaces used in subsonic aircraft control cannot be utilised in a supersonic airliner design since they would cause unacceptable drag.

These considerations had a profound influence on the adoption of a long streamIined fuselage and slender delta wing as the basic Concorde configuration. Proving the validity of this aerodynamic shape required 5,000 hours of studies in subsonic, transonic and supersonic wind-tunnels, supported by large-capacity computers and much research. From this long process, the original design emerged refined and enlarged, but not fundamentally changed.

Ground studies, backed by later flight experience, fully vindicated the earlier decision to avoid any form of variable-geometry wing as a means of achieving the supersonic-subsonic performance compromise. In supersonic military aircraft design, where operational economics is of secondary importance, the "swing wing" is a favoured solution to this problem. In the present state of the art, however, the weight and complexity of the swing-wing hinge mechanism rule it out for commercial operation.

Between the prototype design stage and the final definition of the entry-into-service standard aircraft, various configuration changes were introduced, mainly in the areas of the nose and visor, the wing and the rear fuselage. On Concorde, the visor is the retractable upper section of the nose fuselage. This is lowered for maximum forward visibility at landing and take-off and raised in supersonic flight to streamline the nose and protect the flight deck transparencies against heat and pressure. On the prototypes, the visor was of aluminium alloy construction, with cut outs affording a limited degree of forward vision when the visor was raised. It was never expected that such an arrangement would be acceptable either to certification authorities or to airline pilots, but it was adopted as an interim measure pending further aerodynamic testing and predictable advances in glass technology. The fully glazed visor, introduced on the pre-production aircraft and now standardised, provides excellent flight deck visibility.

In flight, the Concorde wing looks beautiful - and beautifully simple. At closer quarters it can be seen for what it is: a most complex aerodynamic form, with a precisely-calculated degree of camber and taper across the wing structure itself and a combination of droop and twist along the leading edge. Camber, taper, droop, and twist all make their contribution to Concorde's good handling characteristics at low speed without prejudicing its supersonic performance. Much of the later stages of the aerodynamic research programme was devoted to detail refinements in outer camber, the wing tips and the leading edge.

In the fuselage, the major configuration change was the extension and re-shaping of the tail section. Compared with the prototypes, the upper line of the fuselage extends further beyond the fin and now runs almost horizontally, and the lower line sweeps up to meet it at the tail. This change had a significant effect in reducing supersonic drag and provided a bonus in increased fuel capacity.

One of the initial Concorde design decisions was the selection of aluminium alloy as the basic structural material, a decision closely linked to the choice of Mach 2 as the design cruise speed. This policy was ultimately to prove a sound one, but it could be implemented only after a painstaking evaluation of available aluminium alloys. For this purpose, many thousands of specimens were tested for mechanical properties, fatigue strength, and resistance to corrosion.

A new requirement, directly related to supersonic operation, involved testing for resistance to creep, the name that engineers give to the deformation of metal caused by interaction between mechanical loadings and high temperatures. Creep was a phenomenon familiar to the aeroengine designer, but something new to the airframe man. To check possible materials for creep-resistance, samples were submitted to long periods of round-the-clock testing in specially designed, automatically-controlled installations.

Finally, the choice was made of a copper based aluminium alloy, known in Britain as RR58 and in France as AU2GN. This alloy had originally been developed for use in gas turbine blades, but the suppliers were able to give assurances that it could be produced in whatever form - sheet, billet, forging - and to whatever unit size the Concorde design might require.

The many other types of material used in Concorde have all had to be rigorously tested to prove their suitability for application in a supersonic airliner. They include the titanium, stainless steel, and Inconel components used in the engine bays, the glazings in the flight deck and cabin windows, and a variety of plastics, paints, sealants, adhesives, and non-ferrous materials.

Exhaustive tests

With the material selection made, the way was clear to start on the planned programme of structural research. Once again, the facts of supersonic life introduced new complications into the test programme. To have any real value, the laboratory tests needed to reproduce the thermal profile of a typical supersonic flight; the sudden rise in skin temperature during acceleration into the supersonic regime, the heat soak during supersonic cruise, and the sudden cooling of the aircraft surface during deceleration to subsonic speed.

The whole Concorde structural programme culminated in the testing of two complete airframes in vast new laboratories built with this objective in view. One airframe was subjected to static load testing at CEAT, Toulouse, and the second is undergoing fatigue testing in the structures laboratory at the Royal Aircraft Establishment, Farnborough.

Static testing of the airframe at CEAT began in September 1969, the first phase of the programme being the imposition of progressive design loads at room temperatures. When the structure had been cleared in these conditions, the tests were repeated in transient and steady temperature conditions representative of actual in-flight operation. An impression of the scale of the test may be gained from the following statistics: 80 servo-controlled hydraulic jacks impose the test loads; kinetic heating simulation is provided by 35,000 infra-red lamps; 70,000 litres of liquid nitrogen are used for cooling, making possible a reduction in skin temperature from +120°C to -10°C within 15 minutes; and the test instrumentation is capable of recording and processing 8,000 data points every two seconds.

The test programme was successfully completed in 1972 and, as a result, the airframe was cleared for 385,0001b. take-off weight. Since that time, further static testing has cleared the structure to a takeoff weight of 400,0001b.

Static testing proves the integrity of an airframe structure in relation to the numerous transient heavy loadings, aerodynamic and mechanical, to which it will be submitted in flight conditions. It is essential to complement these tests by a thorough appraisal of the fatigue life of the airframe; in other words, its ability to sustain, year in and year out, the regularly repeated cycle of loadings imposed in the course of a normal flight. Fatigue testing of the Concorde airframe is thought to be the most elaborate exercise of its kind ever attempted.

In the RAE laboratory, the fatigue test specimen is encased in a kind of outer "glove", providing an annular duct around the airframe through which hot and cold air is pumped to reproduce the flight temperature cycle. Hot water is used for heating the air and refrigerated ammonia liquid for cooling it. Circulation is by means of five 2,300hp motor-fans. A hundred servo-controlled hydraulic jacks are employed for external loading of the specimen and internal loads - representing cabin pressurisation and air conditioning and fuel movement - are also imposed.

To reduce fatigue test time, internal heating and cooling of the specimen have been provided, and maximum temperatures applied are at a higher level (120°C against 100°C) than those encountered in flight. It had been established in earlier tests that the fatigue effects of a soak at a given temperature for a given period of time can be exactly reproduced in a shorter period at a higher temperature. The effects of a one-hour cycle on the RAE rig are therefore equivalent to those of a typical three-hour flight.

Fatigue testing started in August 1973 and by the end of 1974 the certification requirement of 6,800 cycles had been met. For some years to come, it is planned to complete 7,000 flight cycles yearly which will mean that the test specimen will always have built up at least three times as much fatigue life as the earliest aircraft to enter passenger service.

A variety of other tests has been made on different parts of the structure. For example, development work on the flight deck and cabin glazings included static test to failure, fatigue cycling under realistic temperature conditions, and fail-safe testing in which one element of the glazing has been deliberately failed-while the remainder is under load. Cockpit windows have been subjected to bird impact tests, and the whole structure to hail impact tests. An important feature of the programme was acoustic fatigue testing in France and Britain to establish the resistance of the tail and fin structure to the high jet-noise environment.

Like the aircraft structure, all the Concorde systems had to be designed to operate over the much wider range of temperatures that supersonic operation involves. These systems have been individually tested and developed in specially constructed full scale ground test rigs, many of which are impressive engineering achievements in their own right. There are major test installations for the following systems; hydraulics, electrics, flying controls, fuel management, powerplant, undercarriage, and air conditioning. Use of these rigs enabled many systems design problems to be ironed out before the prototypes flew and thus saved much valuable flight development time.

The hydraulics rig is at Aerospatiale's Blagnac, Toulouse, design centre and is a complete replica of the flying control system with the associated hydraulic and electrical systems. It also incorporates the undercarriage functioning system. Adjoining this rig is the Concorde design flight simulator, one of the most advanced installations of its kind in the world. Provision is made for the hydraulics rig to be connected to the simulator flight deck for testing of the flight control system.

Although the simulator has been used to some extent for flight crew training, it is primarily a design tool. Before flight testing began, it was extensively used for investigation of flying characteristics and studies of control-system response, and it has been linked to Air Traffic Control at Orly airport, Paris, to enable ATC authorities to study techniques of integrating supersonic airliners into the existing operational patterns. At the Filton works of BAC there is a simpler form of flight simulator employed for the study of specific design cases,

Also at Filton is the massive fuel systems test rig, which consists basically of a movable platform on which is mounted a complete reproduction of the aircraft's fuel tank system. During a test cycle, the platform is moved to simulate the attitudes and accelerations that the aircraft will experience in flight, and at the actual fuel temperatures and pressures and rates of climb and descent. Use of this rig enabled modifications of the fuel management system to be introduced early in the programme.

Two full-scale rigs were built for electrical systems testing, one for the generation system and the other for the distribution system. There are other important systems rigs, notably those for the engine air intake system, and for undercarriage, wheels, and brakes. Nobody has ever seriously disputed the makers' claims that when it goes into airline service, Concorde will be the most thoroughly tested airliner in aviation history.

Producing the power

Concorde's cruise speed of 1,350mph is equivalent to the muzzle velocity of a .303 rifle bullet. The objective was to design and build a passenger aircraft capable of maintaining this speed for more than two hours at a time. One of the most important problems facing the designers was therefore to produce a powerplant capable of achieving this level of performance.

Payload represents about seven per cent of the total take-off weight of a supersonic transport whereas a typical modern subsonic airliner can carry about 24 per cent of its take-off weight as payload. If, at entry into service, the thrust of the subsonic engines is one or two per cent below the design estimates, the effects on payload will be adverse but not catastrophic. A similar shortfall in the efficiency of supersonic engines would mean the difference between operating at a profit or at a loss.

So the Concorde powerplant designers started their work in the knowledge that they had virtually no margin for error. They had to get it right, and the "it" that they had to get right was a far more difficult problem than any previously encountered in civil powerplant design. These difficulties stemmed from the fact that the airflow requirements of the engine vary considerably in the subsonic, transonic, and supersonic phases of flight. Use of the term "powerplant" is not just a pompous alternative for the word "engine": it is an overall description of the four components which together produce the motive power; the engine air intake, the engine, the reheat and the exhaust nozzles.

In this vitally important area, airframe and engine manufacturers have worked in the closest collaboration from the beginning. Overall design co-ordination is the responsibility of BAC, who are also responsible for the intake design. The "flange to flange" engine, the Olympus 593 turbojet, is the responsibility of Rolls-Royce Bristol Engine Division, and the nozzle is the responsibility of SNECMA, the French aero-engine firm.

In a gas turbine engine, air is drawn in through an intake and is compressed (and therefore heated) by a compressor driven by a turbine at the rear of the engine. The hot compressed air passes into a combustion chamber where fuel is injected into it the fuel-air mixture is ignited and the hot gases are ejected through the rear jet pipe to provide forward thrust. Between the combustion chamber and the jet pipe, the exhaust gases also drive the turbine.

The Olympus engine

The Olympus used in Concorde is a twinspool turbojet, which is almost equivalent to saying that it is two engines in one, since it has two independent compressors each linked to its own turbine. This design concept was first evolved in Bristol some 25 years ago to meet the engine requirement for the subsonic military aircraft known later as the Vulcan bomber.

Harking back to basic principles, it is clear that if compression of the intake air can be increased, its temperature will rise and less fuel will be needed to produce an equivalent amount of thrust energy. To achieve a higher compression ratio, the Olympus designers adopted the novel solution of using a low pressure and a high pressure compressor running in series.

The Olympus engine is much older than Concorde. In the course of flight development of the first mark of Olympus, a Canberra aircraft, powered with two of the engines, broke the world altitude record as long ago as 1953. The Olympus-powered Vulcan began RAF service in 1956, and the engine thrust was steadily increased from the 11,0001b. of the Mk 101 up to 20,0001b. for the Mk 301.

But the evolution of the Olympus still had a long way to go, and the next step forward was to take it across the border from subsonic to supersonic. A supersonic engine, later designated the Olympus 320, was developed for the BAC TSR2 military aircraft (the initials standing for "Tactical Strike Reconnaissance"), and, although there was no departure from the twin-spool design concept, changes were made in materials to meet the higher operating temperatures.

The first TSR2 made its maiden flight in September 1964, and a second was nearing readiness for flight when, in April 1965, the British government decided to-cancel the project. Britain was left, however, with a technological legacy in the form of a supersonic turbojet engine.

The clinching point when it came to a choice of engine for Europe's supersonic transport was not just the record of the Olympus, but its availability. Other contenders were still at the drawing-board stage, but the Olympus 320 had been built, had run on the test-bed, and had flown first in a Vulcan flying test-bed and then in the TSR2 itself. A detailed account of the subsequent development of the 593 would almost fill a book. Possibly the best way to sum up the remarkable progress that has been made since the inception of the Olympus, is to contrast thrust figures: the Olympus 100 first ran at 9,1401b. thrust and the Olympus 593 has demonstrated a thrust of over 40,0001b. - equal to the total brake horsepower of 10,000 Minis.

One Olympus development must be singled out for special mention. The two prototype Concordes were powered with early versions of the Olympus 593B, which produced much smoke at take-off and landing. To meet the criticism about the exhaust trails Rolls-Royce and SNECMA decided in 1969 to develop for the Olympus a type of combustion chamber which had already demonstrated, in the Viper, Sapphire and Pegasus engines, its ability to eliminate smoke. The introduction of this "annular" type combustion chamber, in conjunction with a new vaporising fuel injector system, has made the Concorde exhaust virtually smoke-free. In this respect Concorde ranks among the cleanest of current aircraft.

Supersonic flight brings. the airframe designer new problems because of the increased structure temperatures it creates; it brings the engine designer new problems because of the higher operating temperatures required to produce the higher thrust. At Mach 2 air will enter the intake at about -60°C, will be compressed in the intake and be at about 130°C when it reaches the face of the engine, and will leave the high-pressure compressor at 550°C.

In order to cope with these extremely high temperatures, materials used in the subsonic Olympus have been superseded. The low-pressure compressor and the first stages of the high-pressure compressor are made in titanium; this not only saves weight but is robust enough to withstand ice, birds and other "foreign objects" that get ingested into the engine. To resist the even higher temperatures further back in the engine, nickel-based alloys are used for the final stage of the high-pressure compressor, the combustion chamber, the turbine blades and the reheat assembly.

There are many criteria by which an aero-engine may be judged. Airline engineering staffs tend to use the yardsticks of reliability and ease of maintenance, and the Olympus designers had these objectives very much in mind from the outset. By the time Concorde enters airline service, its engines will have been more comprehensively tested than any other type and will be backed by the experience of 46,000 hours of operation on test-beds or in flight. The Olympus is also a modular design and can be broken down into 12 main assemblies to speed up overhaul procedures. Internal inspection of the engine can be made, without removing it from its nacelle, by borescope, an instrument first developed for use in aero-engine research and ground testing.

Air intake system

The engine air intake system is one of the most remarkable pieces of equipment in the Concorde, and its efficiency is of critical importance to the overall performance of the aircraft. Taking in air at speeds up to Mach 2.2, the intake has to deliver it in an even flow to the face of the engine at a speed of Mach 0.5. So, at supersonic cruise, there is a four-fold deceleration of the intake air from 1,350mph to about 350mph in the length of the intake, a distance of 11ft. Apart from that, the amount of air (the mass air flow) has to be precisely adjusted to the requirements of the engines which as already explained, vary considerably over the speed range.

These necessary variations in mass airflow can be achieved only by altering the size of the intake throat; by making it, in the engineering term, a variable geometry intake rather than a fixed one. In Concorde the variation of the intake geometry is obtained by the use of two movable ramps in the roof of the intake and a spill door in the intake floor. In the spill door there is a small flap that opens inwards to serve as an auxiliary air inlet when required. Automatic control of the movable ramps is effected through electronic black boxes. Pressure sensors in the intake provide continuous information from which the control unit is able to calculate the right position for the ramps and actuate the ramp mechanism accordingly.

At take-off and in climb the ramps are in the fully open position, the spill door is closed and its inlet flap is open (the intake configuration at landing is similar). As subsonic speed increases in the climb, the flap is gradually closed but the ramps remain open until the aircraft reaches a speed of about Mach 1.3 (nearly 1,000mph). As supersonic speed increases above that point, the ramps are automatically lowered and this sets up a series of controlled shock waves within the intake. These shock waves slow the intake air down to subsonic speed before it reaches the engine.

When Concorde's powerplant was designed, it was regarded as a daring innovation to propose the use of reheat on an engine for a civil aircraft. Up till that time, reheat systems had been used only on military aeroplanes. The reheat system is situated aft of the turbine and is used to provide additional thrust by igniting fuel in the jet pipe.

For a supersonic transport, one advantage of reheat is that it provides a considerable increase in take-off thrust without any great weight penalty. Reheat is also employed in transonic acceleration to reduce the total flight fuel consumption, for although the use of reheat during the acceleration phase increases the immediate fuel consumption, there is a benefit in getting faster to the supersonic cruise speed at which consumption is at its lowest.

Exhaust nozzles

In a supersonic aircraft, the exhaust system has to incorporate variable geometry features to provide the required variations in the exhaust gas stream. This is the one area of the Concorde powerplant that has undergone major redevelopment since the original design. Subsonic aircraft exhaust their turbojet gases through a convergent nozzle which, by forcing the gases to pass through a smaller orifice, increase the thrust they produce. Engine efficiency in a supersonic aircraft can be improved by installing a second divergent nozzle to permit expansion of the gases. A convergent-divergent exhaust nozzle system was an integral part of the Concorde powerplant design from the beginning, and its development to the present high level of efficiency is a considerable achievement.

One of the many problems has been to provide a divergent nozzle that allowed the required exhaust expansion without causing external drag in supersonic operation. Both the exhaust nozzles are variable geometry features. Variations in the area of the primary nozzle are used to control the turbine temperatures and rpm while variations in the secondary nozzle area ensure the efficient expansion of the primary gas stream

The production-standard secondary exhaust nozzle, known as the TRA (for "Thrust Reverser Aft"), is notable as an area of Concorde in which American designers have made a direct contribution. In this nozzle the main components are two "buckets," shaped like clamshells, which open and close vertically across the jet streams from each pair of engines. To meet the severe operating conditions, the buckets are made of a special welded steel honeycomb material whose manufacturers, Stresskin Products Co. of California worked closely with SNECMA and BAC engineers on the design of the TRA nozzle.

At take- off the secondary nozzle is partially closed, and during this phase the movement of the exhaust buckets, in conjunction with operation of the silencers, reduces the take-off noise levels. During the transonic acceleration, the buckets are moved progressively to the fully open position which they maintain throughout supersonic cruise. For reverse thrust, the buckets are closed across the gas stream, thus deflecting it in a forward direction upwards and downwards.

Deciding on a maximum speed

In any supersonic transport design, the most critical decision is the choice of cruise speed, for this has a profound bearing on the aircraft's aerodynamic configuration, its structure and the type of material used. Working independently, French and British engineers had opted for a Mach 2.2 (1,400mph) cruise speed. Later in the Concorde project this was slightly reduced, for structural life reasons, to Mach 2 (1,350mph). There are two basic reasons why the French and British - and the Russians - thought that Mach 2.2 or thereabouts was the right speed for supersonic air travel. One, variation of structure temperatures with speed; and two, overall efficiency.

Supersonic aircraft fly slower than sound as well as faster than sound. Therefore the Concorde structure and all its systems had to be designed to function efficiently over a much wider temperature range than those of any previous airliner. They had to cope with temperatures as low as-45°C, encountered flying at subsonic speeds in the icy upper atmosphere, to + 150°C, encountered in supersonic flight, for the temperature of an aircraft structure rises rapidly as the cruise speed increases.

At Mach 1 (750mph) the average temperature of the structure is still just below 0°C, but at Mach 2 it has risen to about 120°C, higher than the boiling point of water. By the time a cruise speed of Mach 3 (2,000mph) is reached, the structure temperature has risen to about 300°C, approaching the melting point of lead and well above that of tin! Materials testing was already under way in both countries, and the indications were that the upper temperature limit that could be accepted for an aluminium structure was around Mach 2.2. To go for anything much faster than that would mean designing the structure in a stainless steel or titanium alloy. This would have meant a much longer development period and greatly increased development costs.

By that time, aircraft designers and production engineers had been specifying and using aluminium alloys for over 40 -years and were thoroughly familiar with its virtues and its limitations. That was one powerful motive for the decision to build the SST in aluminium. Another was the minimal time advantage that a Mach 3 airliner would show over a Mach 2.2 type. A Mach 2.2 airliner would halve the subsonic flight time across the North Atlantic from seven hours to three and a half hours but the Mach 3 airliner would cut only an additional 20 minutes of the Mach 2.2 time.

Two factors had to be taken into account in assessing overall efficiency - aerodynamic efficiency and propulsive efficiency. Aerodynamic efficiency is at its best at a speed of Mach 0.80 (600mph). As speed rises to Mach 1 (750mph) there is a sudden falling off. The sharp decline continues until about Mach 1.4 (1,050mph), and, between Mach 1.4 and Mach 3 (2,000mph), although there is still a decline, it is much more gradual.

This decline in aerodynamic efficiency is offset, however, by the propulsive efficiency "curve" of the turbojet engine, which rises steadily to Mach 3 and a little beyond. (It should be made clear at this point that we are speaking of the straight turbojet: the efficiency of the high bypass-ratio engine - the fanjet used in modern wide-bodied subsonic airliners - starts to decline just before the speed of sound is reached).

The result of offsetting the loss of aerodynamic efficiency against the increased propulsive efficiency of the turbojet as speed increases, is that overall efficiency has recovered to something near subsonic level at Mach 2, and that from Mach 2 to Mach 3 there is a gradual improvement. In the Anglo-French view, however, the advantages of going beyond Mach 2.2 were not worth the development time and costs it would require.

All this may seem obvious and clear cut now, but it was not so at the time. These European views were strongly challenged by eminent SST engineers in the USA who favoured a Mach 3 speed, and it took a great deal of confidence on the part of French and British designers to uphold their convictions. Many people in Europe accepted the American Mach 3 thesis, including some of the more vocal critics of Concorde in Britain and France, who argued that Concorde was too small and too slow, and would soon be outmoded by the bigger, faster American SST.

Getting the wing right

Concorde's final wing shape is the outcome of a number of design compromises. The basic slender delta is a good shape for supersonic flight because it can be designed and built for low drag at high speeds, but it is far from being the ideal shape for sub sonic speeds. Good control and handling qualities have to be "designed into" the slender delta.

In the Concorde design, the option of using a variable-geometry wing - a swing-wing - in order to get the best of both supersonic and subsonic worlds had been firmly rejected on grounds of weight and complexity. It took a continuing programme of research and development extending over a number of years to evolve the final wing shape. This was necessarily a lengthy process because all the successive changes had to be exhaustively tested, first in the wind-tunnels and then in the air. Increases in wing area and subtle reshaping of the wing tips and leading edges have made for better low-speed controllability and an improved lift-drag ratio at subsonic speeds.

Wing-design was another factor favouring the adoption at Mach 2.2 as the upper limit for the cruise speed. At that speed the designer can achieve a good working compromise between conflicting supersonic and subsonic aerodynamic requirements. As the cruise speed increases beyond Mach 2.2 the wing has to be made more slender to counter the increased drag, but the more slender it becomes the more difficult it is to achieve good low-speed handling.

The Two Centres

There is yet another design problem peculiar to supersonic aircraft - the means used for trimming the aircraft in flight. This involves the aircraft's "aerodynamic centre," the point along the aircraft's length through which the lift forces act, and also the "centre of gravity," which is the point through which the aircraft's weight acts. When you "trim" an aircraft you take action to keep these two centres in the right relationship with each other.

On a subsonic aeroplane, the aerodynamic centre moves only within fairly narrow limits and trimming can be done aerodynamically. Movements of trim tabs (small sections of the flying control surfaces), or of a moving tailplane, will alter the airflow over the control surfaces to make the necessary small changes in the centre of gravity position. In this area, as in many others, the supersonic designer's problem is more complicated.

When a supersonic aircraft accelerates from subsonic speed up through the transonic range to supersonic speed, its aerodynamic centre moves aft. In Concorde, the cambering of the wing and the "wine-glass" line of the leading edge both help to reduce this movement, but it is still appreciable. If no action were taken to change the centre of gravity, the result of moving the centre of lift towards the rear would be to raise the tail end of the aircraft and so put it into a nose-down attitude, resulting in greater drag and more difficult control.

To make the adjustment by aerodynamic methods as in the subsonic aircraft is not feasible because any deflection of flying control surfaces would have to be made throughout the supersonic cruise and would cause unacceptable drag. In Concorde the method used to keep the centre of gravity in the right place is to pump fuel between the main tanks and forward and aft trim tanks as required.

It is true that the supersonic designer is beset with new problems, but the airflow characteristics of the slender delta give him at least one natural advantage over the subsonic designer. Nature has provided the delta wing with its own "high lift device" and there is no need for the complicated arrangement of slots and flaps that can be seen on a subsonic wing when the aircraft takes off or comes in to land.

To make its landing approach at a safe speed, a delta aircraft assumes a fairly steep angle of attack. At take-off it is also at a steeper angle than a subsonic type. At these high angles of incidence, the airflow over the leading edge of the wing breaks away and forms a vortex. If this vortex formation remains stable (and the Concorde leading edge is designed to keep it stable) it follows the line of the leading

edge and has the effect of producing additional lift in just those two phases of flight - landing and take-off- when the subsonic aircraft has to use mechanical means to get added lift.

The droop snoot

Besides the powerplant and fuel system, the Concorde's adjustable droop nose is another design feature that differs markedly from subsonic practice. A movable nose is required on a supersonic airliner to give the pilots a good view of the airport runway in landing and take-off. The nose of a supersonic transport, unlike the blunt rounded front end of the subsonic airliner, is streamlined for high speeds. It is long and tapers to a sharp point. Concorde comes in to land and takes off at a more nose-up attitude than a subsonic aircraft, and, if it were fixed, the long nose could hamper the view from the flight deck on to the runway.

To get over this problem, the nose unit (all that part of the fuselage forward of the flight deck) is made so that it can be lowered during landing and take-off and raised during the other phases of the flight. The droop nose is composed of two sections: the main nose structure and the glazed upper section of the forward fuselage, known as the visor, which can be lowered and raised independently.

In supersonic cruise, the nose and the visor are raised. This streamlines the front end of the aircraft to minimise air resistance, and the visor protects the flight deck windows against kinetic heating and air pressure. Forward view from the flight deck is through the flight windows and the visor windows. At take-off and in the early stage of the subsonic climb, the visor is lowered and the nose is lowered to its intermediate position, 5° of droop. In subsonic cruise the visor is lowered but the nose is raised. In the approach, and at landing, and also while taxying at the airport, the visor is lowered and the nose is in the down position, 12.5° of droop.

With the droop nose, which has been thoroughly tested in the most extreme climatic conditions, the Concorde pilot's view on to the runway is better than that from the flight deck of most other airliners.