GB2136183A - Microprocessor Controlled Gas Turbine Aero-Engine with Navigational Landing Aid - Google Patents

Abstract

The two inventions of which this abstract and summary represents, deal exclusively with aircraft safety, with some emphasis on fuel efficiency as a secondary consideration, in so far as efficiency in specific fuel consumption relates to the overall safety of aircraft operation. Since the Second World War, when Radar was first used, to locate aircraft, air-lanes have become saturated with aircraft. Continued efforts on the part of the manufacturers and airline operators have resulted in increased safety, and airworthiness. Despite this catastrophic accidents have occured. Many such accidents have been attributed to pilot error, both on landing and take-off, whilst others have been attributed to misunderstanding between ground control and pilot. Others undoubtedly can be attributed to airframe structural failure, whilst others have been attributed to engine failure in inclement weather especially on take-off and landing. These inventions set out to provide solutions to such hazards, which frequently occur in mountainous country, and during inclement weather, such as blizzards accompanied by sub-zero temperatures, and/or foggy weather, when blind landing devices are unable to cope with such situations without fear of tragedy. Such fear is well founded, due in main to tragic case histories. Space age technology has been used to attack these problems, such as the use of the Zirconium probe to check the combustion efficiency of the jet engine, immediately prior to take-off, and landing. Another product of Space Age Technology has been the radio controlled infrared television camera, which is used in conjunction with the electronically readable photo-electric theodolite. Such a device is described which will give the pilot a visual picture of the topography of the ground as well as an idealised flight path to follow down to the airport, projected onto a VDU and used during inclement weather, which will obviate misunderstanding between the pilot and ground control. Engine failure which is often caused by flame-out conditions resulting from too rich or too lean air/fuel mixtures is solved using the zirconium probe. <IMAGE>

Description

SPECIFICATION Microprocessor Controlled Gas Turbine Aero Engine with Navigational Landing Aid The first invention relates to the improvement in safety and efficiency in gas turbine engines which may have a variable number of combustion burner nozzles in use, in order to maintain stoichiometric combustions of the fuel and oxygen in the air, within the burner flame, in the combustion chambers, and in the reheat unit where fitted. The process is supervised by a microprocessor and zirconium probe. Another distinguishing feature of this process is that pure air, uncontaminated with nitrous oxides is ducted from the H.P. compressor by way of the hollow drive shaft to the reheat burners.

The second invention relates to an improvement to the safety of aircraft in landing in inclement weather, in which the microprocessor is used in conjunction with a photo-theodolite infra-red camera, which maps the exact position of the aircraft in relation to the runway and the contours of the surrounding countryside.

In particular there are two types of gas turbine to which the invention is applicable. The first type is the turbojet, in which the microprocessor would be hard wire programmed to fly the aircraft under all prevailing ambient conditions within maximum efficiency parameters, in which it would also control and adjust the pressure drop across the turbine so as to remain constant. It will perform this function by adjusting the air inlet vanes, and jet pipe propelling nozzle concurrently, unlike present systems which perform this function in steps. The second type of gas turbine engine, which is considered in this application for Patent protection is the turbo-propeller, in which for maximum fuel economy the microprocessor would be programmed to synchronise the degree of propeller pitch and the jet propelling nozzle area for the most efficient use of the engine power.In both types of engine there is a variable number of burner nozzles in-use which would be related to the forward speed of the aircraft, ambient conditions, and engine constraints imposed by the engine design which are: (1 ) temperature of turbine blades, including thermal shock.

(2) temperature of jet pipe.

(3) over-spinning of low pressure compressor.

(4) temperature of combustion chambers.

(5) limiting pressures on jet pipe and all other components.

(6) Avoidance of flame-out, or engine fire conditions at all times.

The requirements for efficient jet propulsion are: (a) Stoichiometric combustion of fuel-air mixtures at the burner nozzles.

(b)- Precise acceleration of the exhaust gases to exit through the variable area adjustable propelling nozzle at ambient atmospheric pressure.

(c) Use of minimum weight of air for cooling the combustion chamber walls, compatible with maintaining a constant temperature within the design permissible temperature, which also applies to the turbine blades.

(d) In the case of the turbo-prop engine, feathering the propeller to maintain efficient propulsion, that is-maintaining the correct ratio of propeller thrust to jet thrust with respect to forward speed and ambient pressure.

(e) Maintaining high mean blade speed of the turbine since the gas pressure drop is proportional to the square of the blade speed, through the turbine. For a given output the gas velocities, deflections and hence losses, are reduced with high mean blade speed.

(f) The gas jet stream should be uniform and free from eddying and rotational motion, and excessive residual whirl reduces the efficiency of the exhaust system.

(g) Maximum power to weight ratio for all engine components.

Thus an autonomous control circuit for the turbo-jet or turbo-prop engine will be the microprocessor, comprising the following components: (I) Memory mapped visual display unit (VDU), fitted with a comparator, line terminator, and two wire data acquisition, and transmission system which would continously scan the various engine parameters, and compare them with the engine constraint parameters, including the oxygen content of the reference burner exhaust gases as measured by the zirconium probe.

(II) Microprocessor unit (MPU), to analyse the acquisition data and adjust the number of burners in use, in order to maintain stoichiometric combustion at all times and under all ambient conditions, as well as adjusting the degree of propeller pitch in the case of the turbo-prop jet engine, to maintain the most efficient ratio of propeller thrust to jet stream thrust for all air speeds.

(III) Read only Memory (ROM), to store the information which is being processed.

(IV) Random Access Memory (RAM), to store the hard wird operating programme.

(V) As an alternative to ROM, the Microprocessor may be fitted with an Electrically Programmable Read Only Memory (EPROM) which is programmed using a high voltage (25V) and Emulator, and erased using a source of ultra violet (UV) light.

(VI) Input/Output device to interface with the MPU, and the valves, solenoids, servo motors, limiting switches, via suitable bufferi-ng and amplifiers.

(VII) Clock generator, address decoder, address bus, data bus, and line terminator.

(VIII) Zirconium probe to measure the oxygen content in the exhaust gases and provide a reference signal (RS) to adjust the air/fuel of the main engine burners by sampling the exhaust gases of the mimic reference burner.

(IX) Transducers to measure ambient air pressure, L.P. Compressor pressure, H.P.

Compressor pressure, propelling nozzle exhaust pressure, and the pressure drop across the turbine blades.

(X) transducers to measure the output pressure of the high and low pressure fuel pumps, and with transducers fitted to the leading edge of the wing, to measure the slipstream pressure the aircraft airspeed may be deduced.

(Xl) Thermocouples to measure, record, and adjust the temperature of the turbine inlet temperatures, exhaust gases, ambient inlet temperature, high pressure compressor temperature, and combustion chamber tem pe ratu res.

(XII) Thermocouples to measure the sample gas fed to the zirconium probe which will be cooled, and kept below the temperature of the zirconium probe which is heated, and will also require a thermocouple to measure its temperature. Thermocouples may also be used to record the engine bearing temperatures, and other components of the engine which are critical to its operational life.

(XIII) Digital revolution counter to record and control the engine rotational speed, or in the case of an engine comprising a number of shafts, the rotational speed of each shaft will be monitored to a suitable time base and on high by-pass ratio engines that features an intermediate compressor system, between the high and low pressure turbines.

(XIV) An input facility in the form of a key board and memory.

Stoichiometric combustion is defined as the ratio of fuel to air by weight in which all the fuel and all the oxygen in the air at the burner nozzle is burnt leaving no losses due to unburnt fuel, or losses due to excess air, and gives 100% combustion efficiency. The installation of the Zirconium probe in gas turbine engines would require a reference burner of low heat output, from which all the products of combustion would be ducted via a small cooler to cool the exhaust gas stream to be analysed. It also requires a controlled heater supply to keep the probe temperature above the process gas temperature and contains a thermocouple to sense the heater temperature. The sensing probe produces a mV signal logarithmically proportional to the clean dry instrument air which is required to provide the zirconium cell with a reference level of oxygen (i.e.; 20.95% oxygen).The probe continuously monitors the oxygen level of the gases and has a fast response to oxygen changes which is the ideal requirement of closed loop control.

Microprocessor control of the engine to achieve stoichiometric combustion under varying ambient conditions such as those encountered by aero-engines, is most conveniently performed by utilising a large number of burner nozzles arranged in clusters in order to avoid ignition problems, and accordingly there would be clusters of five in the main combustion chambers and reheat section, though the actual number chosen will depend on the versatility and performance required of the engine as well as its ability to cope with varying ambient conditions, without fuel enrichment above that required for stoichiometric combustion or, conversely burning too lean mixtures which chill the flame and can lead to flame-out conditions to pervade the combustion chambers.Such losses which can occur with state-of-the-art engines is shown in Fig. (1) as well as the area of maximum combustion efficiency within which this control system shall operate.

The following reference numbers apply to Fig.

(1) 1. Percentage heat loss in exhaust gases.

2. Percentage of total air in burner nozzle.

3. Losses due to unburnt fuel.

4. Losses due to excess air.

5. Area of maximum combustion efficiency.

With present day aero-engines, the usual method of varying the fuel flow to a fixed number of burners is by adjusting the output of the high pressure (HP) fuel pump. This is effected through a servo-system in response to some or all of the following: 1) Throttle movement.

2) Air temperature and pressure.

3) Rapid acceleration and deceleration.

4) Signals of engine speed, engine gas temperature, and compressor delivery pressure.

With the proposed system these signals would be received by the microprocessor in addition to the correct ratio of air to fuel which may vary with ambient conditions, as determined by the air-fuel reference burner and Zirconium probe, and translate the input data into an output signal which will adjust the air-fuel ratio by adjusting the output of the HP pump, as well as the number of combustion burner nozzles in use, such that stoichiometric combustion is achieved under all ambient conditions and at all aircraft airspeeds.The main advantage of varying the number of burner nozzles in-flight would mean that the fuel flow rate at delivery to the nozzles would vary less than in present day systems, and in addition it would be much easier to achieve good delivery of stoichiometric mixtures without the disadvantage of lag periods during which the mixture is too rich prior to the air pressure building up during periods of increasing engine thrust. Present day systems also suffer the deficiency of relying on mechanical type devices for control which can be slow to respond, as well as out of calibration, and consequentially inaccurate, without the means of continuously checking their own accuracy, which is one of the characteristics of the proposed probe, which can continuously check and correct its own accuracy as an instrument.

Atmospheric pollution is caused by present day systems, and is particularly bad where reheat units are used. Such systems are characterised by emissions of clouds of black smoke which is caused by excessive fuel enrichment accompanied by oxygen starvation. Such systems rely on oxygen present in the exhaust gases, unlike the proposed system which will burn pure air ducted from the H.P. compressor to the nozzle burner via the hollow driveshaft. For this reason and that of noise abatement many large by-pass engines used for passenger travel do not use reheat units.On the other hand reheat units can be used to improve power to weight ratios, as well as enhancing overall performance, of such aircraft, and as a standby power plant for improved take-off from short runway airports such as Washington DC or for take-off from such airports that require a steep climb to clear the surrounding mountains such as at Rio de Janeiro.

Again the continuous use of some reheat burners would improve the specific fuel consumption s.f.c.

if related to the forward speed of the aircraft, and it can be shown that for high forward speeds the discreet use of aterburning can increase the thrust sevenfold. Reheat units as the name implies increases the velocity of the exhaust gases by adding heat, and could be used to particular advantage with high by-pass engines in which one of the characteristics is the low temperature of the exiting gas stream. Furthermore since safety is of paramount importance in the operation of civil avation aircraft such a unit must enhance the safety of any aircraft.Thus with a variable thrust stoichiometric burning engine under the control of a microprocessor which is programmed to run the engine within maximum efficiency parameters, that for certain forward speeds, and relevant mass flow of ambient air, it becomes a practical proposition to use some reheat burner nozzles to improve s.f.c. as well as overall efficiency.

Without infringing the integrity of the present proportional air-fuel control system, it is proposed to add dn autonomous microprocessor controlled circuit which would come into operation under an automatic throttle. It would be known as a trimming circuit, and its main function would entail the continuous monitoring of ambient pressure, temperature and humidity, adjust the air-fuel ratio so that stoichiometric combustion is maintained for all aircraft speeds, maintain a constant pressure drop across the turbine, by synchronously adjusting the air inlet vanes which controls the quantity of air entering the combustion chambers, with the area of the jet pipe propelling nozzle so that the exhaust gases exit at approximately atmospheric ambient pressure.With respect to the latter parameter for efficient jet propulsion, it must be stated that with supersonic streams of gases, where the exit gas pressure is greater or less than atmospheric pressure, the nozzle is said to either under- or over-expanded -respectively. Slightly underexpansion causes negligible reduction in thrust.

With respect to point (c) (see page 2), it is clear that for efficient propulsion that the less air which is compressed for the purpose of cooling the flame tubes, the more efficient the engine will function, since less work will be expended merely to provide cooling. Consideration should therefore be given to improving the relecting qualtities as well as the heat resisting qualities of the flame tube surfaces, with the use of a ceramic material such as Zirconia, which can withstand temperatures in excess of 2,0000C and is better at adhereing to the surface of the coated metal than most, since it propogates minute cracks over its surface which allows for inequalities in rates of expansion between the coated metal and the zirconia.With all ceramic surfacing it is important to match the co-efficients of expansion, and to use a ceramic material which has a very low thermal condutivity co-efficient, which is one of the qualities of Zerconia.

The design of the hard wire programme embodying the maximum efficiency parameters applicable under all ambient conditions, and aircraft speeds can best be achieved, and performed by computer analyses, and synthesis, which can simulate various in-flight ambient conditions and match various combinations of combustion chamber and reheat burners to provide the most economic thrust, with respect to forward speed, or in the case of the turbo-prop the feathering of the propeller, and jet thrust with respect to forward speed. The main difference between the propeller and the jet, apart from their conversion efficiencies, is that the former exerts a thrust by imparting a small increase in velocity to a relative large mass of air, whereas the jet discharges a much smaller mass flow at a much higher velocity.There is thus an efficiency parameter which the microprocessor programme would be programmed to maintain, which is the thrust of the propeller, to the jet thrust, to the forward speed of the aircraft, as well as maintaining stoichiometric combustion under all ambient conditions, and the amount of thrust from the propeller is determined from the pitch of the propeller which is adjustable and known as feathering.

The turbo-prop in which the jet velocity is reduced by using some of its energy to drive the propeller from the turbine is most efficient over a limited range of sub-sonic speeds, but with this efficiency parameter under the control of the microprocessor not only would the efficiency of present-day engines be improved, but the operating range of speeds up to high sub-sonic would be greatly improved. For each type of engine there will be a characteristic speed for different ambient conditions at which the engine will be most efficient with respect to over-all fuel economy, and for engines designed using this invention this speed would be used for cruising, and written into the engine specification for the use of the operator.To facilitate the pilot to fly the aircraft at the most economic airspeeds, the memory mapped VDU would have a colour display in which red would be used for displaying the aircraft airspeed whenever it was flying outside the maximum efficiency parameters, and green when it was flying within them.

The main advantage in using a memory mapped visual display unit is readily understood if one considers that the vast amount of engine data which could be data logged with such an engine management system and need not be displayed on the screen, but only those details requested by the operator, though engine constraints which have reached a critical value would automatically be flashed on the screen in red characters. It is also economic in that the memory space usually used for projecting the data onto the screen can be used for data storage, and with the use of a comparator it need only project that data which is required or critical to the safe operation of the engine. The comparator would be an integral part of the microprocessor which would contain all the critical values of the safe parameters within which the engine must operate.

In order to facilitate the writing of the microprocessor programme for the trimming circuit, it is necessary to obtain test bed results for the diagrammatic graphs shown on Figs. 2, and Fig. 3, in which the following reference letters denote: U Forward speed of aircraft.

T Total number of burner nozzles in use.

S Stoichiometric combustion curve for maximum combustion efficiency.

N Number of Combustion chamber burner nozzles for use.

R Ratio of reheat chamber burner nozzles in use to total number of burner nozzles in use.

E Froude's engine efficiency.

m.e.c. Maximum efficiency curve with respect to specific fuel consumption.

A.L.F. Aircraft level flight.

Fig. 3 approximates to the aircraft flight envelope.

Nr Number of reheat burners in use.

Fig. 4 applies to turbo prop engines to which the following reference letters denote: P 1 denotes ratio of aircraft propulsion from propeller.

P2 denotes ratio of aircraft propulsion from jet stream.

S Maximum efficiency curve (idealised).

U Forward speed of aircraft.

Direction of arrow indicates increasing values.

In order to facilitate the writing of the microprocessor programme which embodies the efficiency parameters, it is necessary to obtain test bed results for the graphs shown in Figs. (2 to 4) inclusive, and sufficient graphs must be plotted to cover various combinations of ambient temperature, pressure, humidity and forward speed of the proposed aircraft. These variations should be limited to the maximum ceiling height to which the aircraft is proposed to fly, whilst the temperatures and humidity will be determined from variations shown on world maps. Forward speeds must embrace variations caused by winds and gusts up to force 12 as measured on the Beaufort Scale.

Such a gas turbine design must embody a new burner arrangement, and such burners will be fed the exact weight of fuel which will be injected with the air into the combustion chamber where all combustion of the fuel and oxygen takes place at the burner nozzle within the flame tulip formation. It will comprise a spray nozzle type burner, which essentially comprises a fuel inlet port, fuel strainer, feed arm, air inlet port, swirl slot, and deflector cone.

Clearly such a multi-burner nozzle arrangement which burns stoichiometric mixtures, with a variable heat output means that the combustion stability is assured and that the flame will remain alight over a wide operating range, which may include gliding, diving, and engine idling, which unlike state-of-the-art engines is sometimes extinguished, due to a high airflow, and only a small fuel flow, i.e.; a very weak mixture. Again with present-day engines, the range of air/fuel ratios between the rich and weak limits is reduced with an increase in air velocity, and if the air mass flow is increased beyond a certain limit flame extinction occurs.This phenomenon gives rise to a combustion stability loop, which is simply a graph of air/fuel ratios plotted against air mass flow Kg/sec within which the engine must operate, and covers the required air/fuel ratios and mass flow of the combustion chamber. With the proposed method of restricting the number of burners in use, and feeding the burners directly with stoichiometric mixtures, no such problems arise, and flame extinction is avoided, as well as atmospheric pollution from over-rich mixtures.

The lowest limitation of power output is determined from the least number of permissible burners required to maintain the minimum rotational speed, or in the case of the turbo-prop the minimum rotational speed of the propeller, without causing engine stalling.

Again present-day engines, which operate as high pressure ratios engines produce exhaust smoke at take-off, caused by carbon particles being formed in over-rich regions of the primary combustion zone in conditions of low turbulence, high pressure, and temperature, at the burner nozzle, and this cannot occur with the proposed system of stoichiometric combustion, neither will there be ignition problems from weak or rich limits. It will also be possible for the microprocessor to establish the best fuel to air ratio for instant ignition during freezing cold conditions, when it is more difficult to establish combustion under 'cold' conditions than to maintain normal burning.

The microprocessor will monitor fuel consumption during flight by providing a continuous update on computed reserves forecast to be available on arrival at the destination, and where these are critical making recommendations on safe air-speed, in order to maintain safe reserves, and flashing on the screen such information.

Such an engine as the 'microprocessor controlled gas turbine engine' M.C.G.T.E. must also control water injection with respect to ambient air conditions, and with respect to metering the quantity, for which it is ideally suited. It must therefore be restated that the maximum power of the gas turbine engine depends to a large extent upon the density of weight of airflow passing through the engine.

There is thus a reduction in thrust or shaft torque as the atmospheric pressure decreases with altitude, and/or the ambient temperature of the air increases. Under such conditions there will be an exact weight and rate of injection of water and methanol, which will restore the power output, and in some cases (as determined by the computer) boost the power output efficiently for take-off by cooling the airflow with water, or water/methanol. When methanol is added to water it provides anti-freezing qualities, as well as an added source of fuel.

Although present-day engines use two basic methods of injecting the coolant into the airflow, i.e.; spraying the coolant directly into the compressor inlet, or spraying the coolant into the combustion chamber, which is more suitable for axial flow compressor engines, the M.C.G.T.E; will only use the former method. The reason for this is easily understood if one recognises that the integrity of the combustion process is based upon the gas samples taken from the reference burner, and it will also determine from the sample gas whenever it is more efficient to use water injection together with the quantity for any particular ambient conditions in which the engine must operate, and especially at take-off from short runway airports, or where extra boost is required at high altitude.Accordingly, it is proposed to provide an input facility for the microprocessor which will inform the computer section of the microprocessor before attempting take-off from short runway airports in inclement weather, or where extra boost is required during steep climbing required to clear surrounding mountains.

With high altitude flying where the ambient temperature reduces to -56.50 deg C; at 36,089 ft and remains constant up to 65,600 ft cooling the engine is not a big problem. Between these heights ambient pressure reduces from 3.28 Ib/sq in to 0.82 Ib/sq in which effects the engine in two ways.

The fall in pressure reduces the air density and hence the mass flow of air entering the engine for a constant engine speed, which causes the thrust or shaft horsepower to fall, and the proportional fuel control pump adjusts the fuel control rate to match the reduced mass airflow, thus maintaining constant engine speed.

The fall in air temperature, on the other hand increases the density of the air albeit at a lower rate than the fall, in temperature and so compensates to some extent for the loss of thrust.

Between the altitudes mentioned above, where the temperature remains constant the thrust or s.h.p. is effected by pressure only.

High altitude flying may thus be improved using stoichiometroc combustion, by utilising the cluster burner arrangement in which all the combustion air is fed to the cluster in a single tube from the H.P. Compressor and all the air may be fed to a single nozzle burner, in order to sustain stoichiometric combustion within the flame tulip, and would not rely on surplus oxygen in the cooling air to evaporate the fuel sometimes without burning it completely, which occurs with present fuel systems.

The cluster of burners would comprise a primary burner in the middle of the cluster in which the rate of flow and pressure of the fuel is controlled by the throttle setting, and it would be fitted with an air inlet port through which is admitted the correct weight of air for stoichiometric combustion to take place within the flame tulip. In order to maintain the correct mixture ratio, the rate and pressure of the fuel delivered to the burner must always be proportional to the rate and pressure of delivery of the air delivered to the burner. Thus the spray nozzle type burner shown in Fig. 6, is fitted with a pressurizing valve to the air inlet. As the fuel flow and pressure increase, the pressurizing valve moves in direct proportion to admit more air to the burner nozzle. When the fuel stops flowing the air port closes shutting off the air supply.

This type of spray nozzle burner differs from other burners in that all the combustion air is injected with the fuel, and not just a small proportion of it which is the case with present-day spray nozzle burners. When the primary burner has reached its maximum stoichiometric combustion heat out put, and where there are main burners, these will be progressively turnedon by the microprocessor until the cluster reaches its maximum heat output. Not all systems will have, or require main burners for the system to operate successfully, but where they are in use for example on high flying combat aircraft, each main burner will be fitted with an individual pressurizing valve which opens at relatively higher pressures than the primary burner, thus turningon the burners at varying pressures relative to the pressure in the main burners feed tube.For low fuel pressure in the main fuel tube the lowest heat output burner will be actuated, and as the pressure builds-up the next higher heat output burner will be actuated, and so on. This feature is important when flying at high altitude, since efficient atomization of the fuel is also required at low fuel flows, when the primary burner will not be in use, which is the case of altitude, and by aerating the spray, with the correct ratio of air, the local fuel-rich concentrations produced by other types of burner are avoided, reducing both carbon formation, and exhaust smoke. The advantage of using the spray type nozzle is that low pressures give effective atomisation of the fuel thus permitting the use of the comparatively lighter type gear-pump, compared with the plunger type pump.

The stoichiometric burner will now be described with the aid of Fig. 5, to which the following reference numbers apply: 6. Deflector.

7. Swirl slot for atomising fuel.

8. Pre-mixing air inlet port from compressor.

9. Cylinder containing pressurising valve which opens as pressure of fuel increases thus admitting more air, the valve comprises a helical spring and poppet valve, which also acts as a non-return valve.

10. Air outlet from hollow tube.

11. Fuel strainer.

12. Feed arm.

13. Fuel inlet from throttle setting, which determines fuel pressure.

14. Flame protection perforated plate, as well as providing instant evaporation of atomised fuel, to be located, and of such configuration to provide maximum efficiency and protection, under conditions of low output.

1 5. Air inlet from HP compressor, air filter, pulse damper to provide steady state air pressure free from compressor surge pulsations.

1 6. High pressure compressor air flow.

1 7. Lever or push rod to open and close air inlet valve.

18. Piston and cylinder valve which opens and closses air inlet valve in response to fuel pressure, and/or microprocessor controlled signals.

1 9. Atomised air-fuel outlet stream.

20. Air outlet holes from hollow tube to swirl chamber.

The first preferred means of actuation entails replacing the piston valve 18, by an electrical interface, such as a servo motor with drive shaft and worm gearing, or by a solenoid, which in conjunction with an input signal from the microprocessor, compresses the pressurising spring of the air inlet valve, to admit the correct weight of air, by mass flow, in relation to the fuel flow, to provide the burner with the correct stoichiometric ratio of air to fuel for efficient combustion under varying fuel flows. In practice the zirconium probe will provide the MP with a reference signal, which will then be amplified to an adequate force level for this purpose. Such a process which employs a microprocessor closed loop design, and self correcting instrument is scientifically accurate.

By carefully sizing the air inlet valve for the maximum heat output condition any small deviations from stoichiometric air-fuel ratios will automatically be corrected by the Microprocessor acting in conjunction with the air control valve, and zirconium probe reference burner.

With this system and burner arrangement, under the control of the MP is such that stoichiometric combustion at the burner nozzle can be maintained under all ambient conditions. The first condition to consider is at take-off when the engine is stationary, but being spun-up to maximum static thrust. The primary burners are in a state of increasing heat output until at some point any increase in fuel pressure will mean that the burner is being fed too much fuel which will mean that the burner moves outside the stoichiometric combustion ratio as measured by the referenced burner and probe. At this point the first main burner is turned-on by the microprocessor which opens the solenoid fuel line valve.As the fuel pressure continues to rise the three remaining valves are opened at intervals of time long enough to allow the turbine blades to adapt to the increasing temperature without thermal shock. At some point again determined by the reference burner and microprocessor the engine will have reached its maximum power rating maintaining and burning during the spin-up stoichiometric mixtures, and applies to the reheat burners where fitted in addition to the main combustion chamber burners. When the aircraft has reached its altitude position where it levels out for cruising the microprocessor will sequentially turn-off the main burners as the fuel pressure reduces in direct proportion to the throttle setting, thus maintaining stoichiometric combustion.At higher altitudes where fuel flow must be decreased to prevent overspinning of the low pressure compressor, the microprocessor will turn-off the primary burner, and maintain good stoichiometric mixtures using one or more main burners only.

The final condition of operation is descent and landing the aircraft with maximum safety especially in inclement weather when the flame from the burners may be under continuous attack from freezing air, driving snow or sleet and gale force winds. In such conditions catastrophic results may follow an engine failure caused by flame-out occurring at a critical moment prior to touch-down on the runway. It must be stated that with present-day engines the pilot reduces the engine thrust by reducing the amount of fuel fed to the engine and the ratio of fuel to air becomes a very critical factor especially in inclement weather.With the proposed type of burner this is not the case since stoichiometric mixtures are always used, and in order to reduce engine thrust for the descent, the primary burners are switchedoff automatically, and only the main burners would be in use, and these too, may be progressively switched-off until only one remains alight in each flame tube. The reliability of such a combustion system must therefore be considered to be a big improvement over present day systems, since the ability of the flame to remain alight during descent when the aircraft is virtually gliding is greatly enhanced, since the flame is being fed the correct stoichiometric mixture, and will therefore maintain an efficient as well as intense flame tulip, which will be extremely difficult to extinguish during the worst inclement weather, and represents a distinguishing feature from present-day combustion systems, which merely reduce the air mass flow, to fuel mass flow to between the limits of (25 to 1) and (175 to 1), depending on the total air mass flow through the flame tube, and of course such ratios are much more susceptible to flame-out than the correct ratio.The correct stoichiometric ratio is approximately 15 to 1 for the typical type of fuel used in gas turbine aero-engines, and such a flame will have a very hot flame tulip. Such a combustion process will also enhance the safety of aircraft at take-off when a combustion failure can have catastrophic results, which may occur during inclement weather, when freezing cold air enters the engine, accompanied by gusting winds, where over enrichment of the mixture with fuel will also cause flame-out to occur.

It must now be mentioned that 'blind landings' during inclement weather and poor visibility, is one of the chief factors leading to aircraft disasters. It is therefore proposed to fit an infrared television camera in the nose of the aircraft, which will be used to transmit live pictures, on a closed circuit to the aircraft visual display unit (CRT) and to the airport tracking station, by means of a radio transmitter. In order that the camera may map the centre-line of the runway, it is proposed to mount laser beam light emitting elements at each end of the runway, as well as a number down the centre-line of the runway.The microprocessor will be programmed with the necessary formulae and actual length of the runway, so that the pilot may have a continuous 'read-out' of the aircraft height and direction of heading, in addition to a continuous moving picture of the aircraft glide path with respect to the centre-line of the runway which would be displayed on the screen along the north south axis of the screen, and the aircraft as a moving object in the form of a bright dot on the screen, which may have an artificial horizon which moves relative to the attitude of the aircraft, to enable the pilot to land the aircraft evenly on all wheels of the undercarriage. The microprocessor may also be programmed to project the flight path of the proposed landing and so enable the pilot to take remedial action where necessary prior to 'touch-down'.

Infra-red emitters may comprise of a p-n diode, which occurs in the crystal lamp, and the injection laser. When the p-n junction is biased in the forward direction, the current is made up of holes and electrons flowing in opposite directions.

When these holes and electrons meet in the depletion layer, they recombine, and in the case of germanium and silicon where the forbidden gap is small only thermal energy is released, but where gallium arsenide is used which has a large gap, infra-red radiation is emitted.

In the case where X-rays are used, with restricted use for emergencies, high voltage tubes would be mounted in a safe position above the runway with shielding to prevent rays from striking objects on the ground, and rotating the light beam to sweep the sky so that a continuous ray cannot strike the aircraft.

The main advantages of using this type of device for aircraft to locate themselves with respect to the runway, over radar, are the speed of response, and accuracy of the trajectory. Radar is slow to respond and with landing speeds well above 100 m.p.h. it is too slow for accurate landing in inclement weather. Furthermore, the pilot does not have guidance on his flight path which is essential for safe landings, and must rely on being 'talked down' from ground control, which itself relies on radar. It must therefore be concluded that the proposed method will make a considerable contribution to air safety.

The geometry of such necessary calculations to be performed by the microprocessor are shown in Fig. 6. Applying the principles of stereophotogrammetry which is used in aerial surveying, which may be used in this case by the microprocessor, to continuously locate the aircraft and project its location onto the screen with respect to the runway as well as projecting its height above ground, and displaying its glide path to scale.

Stereophotogrammetry is simply the use of consecutive photo-frames to calculate the coordinates of the aircraft in which the following formulae are applicable to Fig. 6. See page 1 9.

Such a camera will have a photo-theodolite lens, i.e.; it will have cross hairs and all scaling of the runway marker 'lights' will be performed automatically by the microprocessor, which will determine the distance in millimetres of the light source emitter, from the two cross hairs, in both vertical and horizontal directions. For accurate location of the aircraft, the runway will be marked with light emitters at each side of the runway half way along the runway, so that there will be four lights forming a cross.

Since this method locates the aircraft with respect to the level of the runway only, it is proposed to project onto the screen an accurate contour map of the countryside surrounding the airport and using the ordnance survey levels to compute the height of the aircraft above ground level, and give a continuous read-out in the corner of the screen of its height and direction heading.

With the runway oriented in the centre of the screen on the north south axis, the contour map will be likewise oriented relative to the axis of the runway, and the aircraft will be shown as a bright spot moving relative to the fixed runway and contour lines. Mountain peaks may be marked in red for warning. Such a system of navigation over and around airports would be foolproof since the ground control would be able to monitor the flight path, and check with radar location, which is sufficiently responsive and accurate for this purpose. The advantages of such a system are obvious, in that the pilot can land his aircraft in any kind of weather without relying on sometimes primitive instructions in a foreign language from ground control. Such instructions can lead to mistakes and past experience has provided the warning and impetus for a new approach such as the method described.Furthermore misunderstandings between pilot and ground control are eliminated with this process.

Fig. (6) represents a typical aerial stereophotogrammetry survey diagram, from which all the unknowns may be found by comparing similar triangles, and in practice the microprocessor will be programmed to do, and give a continuous read-out of the critical parameters such as height above ground and aircraft direction, though the latter will be apparent from the screen. The following letters symbolize the known properties in Fig. 6.

AB centre-line of runway in the longitudinal direction.

P and Q positions from which consecutive photo-shots are taken by infra-red or X-ray, or Radio telescope, theodolite camera.

A and B represent points of light, or radio wave emissions.

Xb length of runway between points of wave emissions.

xb distance of image of light ray from point B from vertical wire on lens.

xa distance of image of light ray from point A from vertical wire on lens.

x'a and x'b are similar distances as seen from point Q.

f represents the fixed focal length of the camera or telescope.

Xb is length of runway which is also known and would form part of the input data to the microprocessor, whilst xa, xb, xa, xb would be read directly by the microprocessor. By similar triangles the unknown horizontal distances Xa and Xp may now be computed.

Similarly the values of za, zb, z'a and z'b represent the distances of the images from the horizontal wire on the lens which are not shown on the diagram, would be electronically read by the microprocessor, and the distances Yp and Yq can similarly be computed.

The camera would be mounted with rotatable axae similar to a theodolite with two electronically readable vernier scales, one to measure vertical deflection of the telescope, and the other to measure horizontal rotation of the telescope. With this information the microprocessor can calculate the height of the aircraft above the runway, its precise location with respect to the runway, and height above the ground by comparing its position with the contour map of the area, in addition to its heading which will be obvious from the screen.

Referring to Fig. 6, it will be clear that Yq is normal to the runway except when the aircraft is flying vertically above the centre-line of the runway, when it will represent the vertical distance to the runway. By using the deflection angles required to maintain focus on the light emitters, and which are continuously read by the microprocessor and held in buffer memories for continuous use, it is a simple matter to solve the relevant right angled triangle to obtain the true vertical height of the aircraft above the runway, as well as the co-ordinates of the aircraft position, of the aircraft, relative to the runway.

The exact position of the aircraft may then be projected onto the VDU screen as a white spot on the contour map. The contour map would be projected onto the screen from the necessary topographical data stored in the microprocessor floppy disc, and loaded into the R.O.M. prior to any particular airport landing. It should also be stated that unlike the making of a survey contour map, the system only requires a single photoshot, assuming the runway to be level, to determine the height and co-ordinates of the aircraft, as perusal of Fig. 6's similar triangles will confirm.

A less expensive and possibly more ingenious method of displaying the height of the aircraft on the VDU together with displaying the white spot locating the aircraft will embody the use of the scanning radar wave. Radar finds the location of the aircraft, and determines its distance from the time it takes the radio pulse to go there and back.

The sending circuit sends out regular, short pulses using electromagnetic waves only, which travel at the speed of light. We therefore have the hypotenuse side and one included angle of a rt.

angled triangle as well as the compass reading for its direction. Thus it is possible using a ground control computer with radio transmitter to transmit the aircraft position in three dimensions and height above ground in machine language for instant location on the aircraft VDU. From consecutive scans of the aircraft, the heading, and direction of the aircraft, will be apparent, both to the pilot and the ground control. Such a system I will call 'A Microprocessor controlled interactive Radar landing aid', and it would be virtually foolproof, leaving nothing to chance. The use of a photo-electric theodolite television camera may still be used to improve blind landing, since as previously pointed-out radar is too slow to provide precise landing instructions immediately prior to touch-down.As the aircraft continues its descent along its glide path to touch-down the T.V. camera will be in continuous communication with the light emitters and providing the pilot with a continuous 'visual image' of his exact position in relation to runway centre-line, unlike radar 'talkdowns' which involve a certain element of risk in such situations. Once large airliners descend below 50 metres, above the runway, they cannot take any abortive action and must land. The reason for this, is that aircraft require a minimum height above the ground in prder to change direction from descent to ascent. It is therefore of paramount importance that the pilot is provided with the most accurate picture of his position with respect to the runway at heights below 50 m.

Thus with this invention the pilot is not only provided with his exact location with respect to the runway, but is also provided with a continuous optimised touch-down flight path which will be projected onto the screen in the form of a dotted line using a different colour to the white spot which will locate the aircraft. Such a flight path will be analogous to the 'talk-down' instructions which are currently used by ground control personnel, but will be free from human error in that they will be computer simulated and well tested and proven during fine weather conditions. All such landings may of course continue to be monitored by present day methods, which will provide a useful back-up service, and still further reduce pilot errors which from past experience has shown us to be one of the most common sources of aircraft disasters.

In order for the camera to be oriented and pointed in the correct direction to photograph the laser beam emitters, such direction finding instructions may be transmitted from the ground control computer, from its own three dimensional radar location of the aircraft and storing in memory the location of the light emitters it will perform the simple co-ordinate geometry calculations and with a radio controlled direction finder point the camera directly, at the laser beams, whilst orienting the laser beams to strike the target lens. Memory mapped CPU's for use with aircraft have the unique advantage of projecting onto the VDU screen only those parameters which have reached a critical level or limit in the operation of an aero-gas-turbine engine, or may be used to project the exact location of the aircraft as mentioned in this application.For critical parameters to appear in real time on a VDU whenever their value is exceeded, or to be shown at the request of the flight engineer, requires an array of memory locations whose contents are used to determine operational variable limits and project onto the screen their actual value whenever such limits are exceeded. The array is scanned cell by cell, and each cell is compared with the limit in the comparator which forms an integral part of the microprocessor. The rules are applied to the entire pattern simultaneously and the cell's new state must be stored separately, from it's previous state, whilst the remainder is scanned. On memory-mapped displays the actual display memory can be used as the array of cells, thus saving memory, and avoid having to copy the entire array to the screen.The advantage is that the reader does not have to search endlessly to find critical parameters. Typical 'Electronic Engine Control Apparatus' is shown and explained in detail in United States Patent No. 4,309,759 published on January 5th 1982, and under the aforementioned title. The proposed circuitry would differ from that described in this publication in that it would embrace a memory mapped visual display unit.

Further improvements and refinements to the optimised and ideal flight path to the approach to the runway would include flag positions at which may be in-set recommended aircraft height, and ground speed. At the point where the aircraft is lined-up with its nose and camera pointing down the centre-line of the runway, the VDU will then switch to live pictures of the laser beams, but will continue to show the aircraft height above the runway, as well as the centre-line of the runway on the north-south axis. The image of the laser beams should then coincide with the centre line and these will be projected onto the screen, as well as the recommended aircraft ground speed shown in the top R.H. corner, and this should allow for gusting winds, which will prevent aircraft stalling, or overshooting the runway.

Claims (9)

1. A method for improving the operation of a turbo-jet, or turbo-prop, aero-engine, with respect to reliability at take-off, landing, and during flight, especially during inclement weather conditions, maintaining stoichiometric combustion within the flame tulip, in which the combustion efficiency may be checked immediately prior to take-off, landing and during flight, using a microprocessor in conjunction with a zirconium probe and reference burner, of low heat output, from which all the products of combustion are ducted, via a small cooler, and the oxygen level of the exhaust continuously monitored for efficient combustion, which is displayed on a memory mapped visual display unit, for warning in red digits, or for verification in green digits where the ratio is satisfactory, in that all the fuel, is burnt with most of the oxygen in the air, within the flame tulip leaving no trace of either in the exhaust gases.

2. A method according to Claim 1, in which there are at least one primary burner and one or more burners in a cluster, per flame tube, in which each burner nozzle is fed the correct weight of air to fuel for stoichiometric combustion, in which there is a spring loaded air inlet valve for the high pressure compressor which is synchronously opened to admit air to the burner by the pressure exerted by the fuel pipe pressure, such that as the pressure in the fuel pipe to the burner increases the mass air flow to the burner increases, to a point where it reaches its maximum heat output, when the microprocessor actuates the main burners sequentially in order to avoid thermal shock with respect to the turbine blades, until the engine reaches its maximum heat output, and power.

3. A microprocessor controlled gas turbine engine in which pure air is ducted from the high pressure compressor, and fed directly to the reheat burner nozzles, and thereby avoiding air contamination with the non-combustible nitrous oxide NOx compounds present in the flame tube exhaust gases, thus maintaining the integrity of the stoichiometric proportional air-fuel control circuit, under all ambient conditions of flight, and avoiding oxygen starvation, which causes the formation of black clouds of unburnt carbon deposits, which pollute the atmosphere.

4. A microprocessor controlled gas turbine engine according to any one of the previous claims, which may vary its thrust output, or shaft horse power by utilising a variable number of reheat burners in conjunction with a fixed, or variable number of combustion chamber nozzles, in accordance with its engine management programme which relates the number of reheat burners in use with the forward speed of the aircraft, and throttle setting such as for take-off when all burners may be used, whilst adjusting the air inlet guide vanes so that the minimum quantity of air is admitted to the combustion chambers, for cooling purposes compatible with maintaining the flame walls within the maximum design temperature, as well as maintaining the temperature of the turbine blades below the maximum design temperature, and synchrondusly adjusting the variable area propelling nozzle so that the pressure drop across the turbine remains constant, and the exhaust gases exit at approximately ambient pressure.

5. A microprocessor controlled gas turbine engine, of the turbo-prop type in which the thrust from the propeller, and the thrust from the exiting gas streams may be related to the aircraft forward speed and proportioned by feathering the propeller so that the most economic fuel consumption may be maintained during flight, using the microprocessor fuel efficiency parameters which relate all the ambient conditions to match the correct ratio and degree of propeller pitch, to the thrust from the jet exhaust streams.

6. A method for improving the blind landing of aircraft with respect to safety and convenience especially during inclement weather in which the pilot can guide his aircraft to touch down by following its position on a contour map of the airport, and surrounding countryside, which is loaded into the microprocessor on a floppy disc, and projected onto the VDU, in which the aircraft is continuously represented as a white spot which is located continuously from ground control radar, which transmits its position and height above ground by radio, which is received by the microprocessor which projects its position and height above ground onto the contour map, which also shows the safest flight path to follow down to the runway approach for the convenience of the pilot to follow.

7. A method for improving the blind landing of aircraft, with respect to safety and convenience, especially during inclement weather, in which the airport runway is shown to scale on a contour map which is projected onto the VDU screen, in which the aircraft may locate itself, using an infra red television camera with theodolite telescope and electronically readable verniers which measure deflection angles in vertical and horizontal directions, which can photograph infra red light emitters located at each end of the runway, and along the centre-line, and in which the microprocessor can compute the co-ordinates of the aircraft, and height above ground and 'continuously project the position of the aircraft onto the screen with respect to the centre-line of the runway, and idealised flight path to follow down to the runway, and give a continuous read out of the aircraft height above ground level, whilst during approach to the runway, a close up of the runway may be projected onto the screen showing the attitude of the aircraft with respect to the runway, and the horizon, enabling the pilot to make small adjustments for a perfect touch down and landing.

8. An interactive computer controlled system in which the ground control microprocessor interfaces with the aircraft microprocessor providing monitoring and checking data on the aircraft position relative to the runway, and height above the ground, whilst also providing automatic control in orienting the telescope of the T.V.

camera enabling it to photograph the infra-red light emitters, whilst the aircraft is still at a safe height to abort the landing, and interfacing with the aircraft's own computed height and location for correlation and minor correction, should this be necessary, in which the ground control microprocessor transmits information in a computer based language which is instantly correlated with the aircraft's own computed position, and the aircraft transmits its location computed from the camera to ground control, so that there is a continuous interaction between ground control and the aircraft, which may use a close up moving picture of the runway on final descent and touchdown based on its own calculated position, which is continuously checked by ground control monitoring.

9. An interactive computer controlled landing system according to Claim 8, in which an infrared television camera is directed from a ground control radiotransmitter, to photograph the infrared light transmitters, at each end of a known length of runway, using an optical theodolite which has cross wires in which the location of the image of the light emitters is measured electronically from both cross wires, and using a telescope of known focal length, runway length, and vertical and horizontal deflection angles which are electronically read from the vernier scales, may calculate the location of the aircraft from similar triangles in addition to its true height above ground level which may be deduced from the contour map, and give a continuous read-out of both location height above ground, and show on the VDU screen a white moving spot which displays its heading, against a background ordnance survey map projected onto the screen from a floppy disk on which the runway is always shown on the north-south axis.