US2566373A - Fuel control system for turbojet engines - Google Patents

Fuel control system for turbojet engines Download PDF

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US2566373A
US2566373A US640293A US64029346A US2566373A US 2566373 A US2566373 A US 2566373A US 640293 A US640293 A US 640293A US 64029346 A US64029346 A US 64029346A US 2566373 A US2566373 A US 2566373A
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fuel
auxiliary
valve
means
control system
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Edward M Redding
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Edward M Redding
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02KJET-PROPULSION PLANTS
    • F02K3/00Plants including a gas turbine driving a compressor or a ducted fan
    • F02K3/08Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof
    • F02K3/10Plants including a gas turbine driving a compressor or a ducted fan with supplementary heating of the working fluid; Control thereof by after-burners

Description

Sept. 4, 1951 E. M. REDDING 2,565,373

FUEL CONTROL SYSTEM FOR TURBOJET ENGINES Filed Jan. 10, 1946 2 Sheets-Sheet 1 Inventor Attorney Edward M. Redd/fig Sept. 4, 1951 E. M. REDDING FUEL CONTROL. SYSTEM FOR TURBOJET ENGINES Filed Jan. 10, 1946 2 Sheets-Sheet 2 Edward M. Re dd/ng Patented Sept. 4, 1951 FUEL CONTROL SYSTEM FOR TURBOJET ENGINES Edward MtRedding, United States hlavy Application January 10, 1946, Serial No. 640,293

1 Claims. (cl. so-ass) (Granted under the act of March 3, 1883,15

This invention relates to aircraft propulsion motors of the reaction type and in particular to their fuel systems.

An object of the invention is to deliver auxiliary fuel to a reaction motor so that a desired optimum exhaust temperature may be maintained irrespective of variations in flight conditions. Another object is to utilize a single fuel governor control to supply fuel to both the mainand auxiliary fuel systems of a reaction motor "so that the respective rates of flow of the latter two will be under control of the former. I

Further objects, advantages and salient feat tures of the inventions will become apparent from the following description, accompanying draw ings and appended claims.

. In the drawings; Fig. 1 represents one embodiment of the invention and shows a longitudinal section of a generally conventional reaction motor together with a diagrammatic illustration of a a fuel sys- Fig. 2;

Fig. 4 is a similar modification, wherein the orifice is made adjustable; and

amended April 30, 1928; 3'70 0. G. 757') The auxiliary system isessentially a reheat de vice which raises the temperature of gases previously cooled by expansion. It may function in the manner previously described, that is, to' increase the thrust of the unit or may reheat the gasesfor subsequent use in another gas turbine stage as shown e. g. in Fig. 5. It is contemplated. accordingly, that another gas turbine 5' could be used in the gas flow-downstream from the auxiliary fuel supply. It is apparent, also, and contemplated, that more than one auxiliary fuel system could be used where successive gas reheat 1 would be advantageous as e. g., fuel manifolds 9 and 9' with flame holders H and II, ignition 's'ystemsproposed or used prior to this invention have not givenoptimum performance under variable flight conditions because of lack of proper control of the auxiliary fuel. It is apparent that if the auxiliary system supplies less than an optimum amount of fuel, the gas temperature in the exhaust pipe will beless than the materials are capable of withstanding and hence the highest performance of the unit will not be obtained,

Fig. 5 illustrates a modification, wherein the arrangement shown in Fig. 1 is "supplemented by additional turbine and reheat devices.

Referring to Fig. 1, I represents a reaction motor having a compressor 2, main fuel supply manifold 3, combustion chamber 4, gas turbine 5 and tail cone 6. Fuel delivered by the manifold is burned in the combustion chamber by air delivered from the compressor, the products of combustion passing through and drivin the turbine, which in turn drives the compressor. The products of combustion then pass into exhaust pipe I and issue from nozzle 8 providing thrust for the unit. These features are regarded as conventional in the art.

In the exhaust pipe I a manifold 9 is provided which delivers fuel into the gas flow by means of a plurality of jets, its combustion raising the temperature of the exhaust gases, adding to their kinetic energy which increases the thrust of the unit over that which would be obtained by the main fuel system alone. At III is illustrated a suitable ignition device to initiate combustion of the fuel supplied by manifold 9 and II is a flame holder to maintain uniform combustion. These latter two devices are conventional in the art.

whereas, if more than an optimum amount of fuel is introduced the temperatureswill be higher than the materials are capable of withstanding and their useful life will be materially impaired. An optimum fuel flow would be one such that the temperatures are the maximum which the materials can withstand for a certain desired useful life'or that which effects an improved overall thermal cycle.

It has been found that at constant R. P. M. the exhaust pipe temperature, resulting from the addition of auxiliary fuel, can be kept substantially constant under variable flight conditions if the amount of fuel introduced into the auxiliary system bears a fixed relationship or ratio at all times to the main fuel supply. The system herein described accomplishes this desirable result.

At 12 is diagrammatically illustrated a governor which at any particular throttle setting will control the delivery of fuel to a combustion chamber at such a rate that constant turbine speed will be branch to metering device I]. These two devices are constructed in much the same manner as conventional gear pumps and together they constitute a proportional fuel divider 30. They are connected by a shaft I8 50 that they operate in unison, but are not mechanically driven, the rotation of the gears being effected by the flow of fuel through the devices. They are so proportioned in size that their respective delivery rates are in a fixed ratio regardless of total rate of flow. Alternately, they could be made the same size and be suitably connected by gearing such that their respective rates of fiow would be in proportion to their rates of rotation.

Meterin device l6, which would normally be designed to deliver a smaller quantity of fuel than H, due to the normally smaller requirements of the primary combustion chamber. is connected to the main fuel manifold 3 by line iii. The other metering device I1 is connected to the auxiliary fuel manifold 9 by line 20. A valve 2| is provided in line 20 to render the auxiliary supply inoperative when its useis not desired. A line 22 by-passes the metering device It and is provided with a valve 23 which is connected to valve 2| by link 24 so that the two valves operate in unison, and when valve 2| is closed, valve 23 is open. Valves 2| and 23 are not metering valves, but simple cut-off valves linked together for reverse operation. This link may, if desired, be connected to the pilots throttle so that valves 2| and 23 are under control of the throttle only as it is moved to its full throttle position; that is, under normal power requirements only the main fuel system leading to manifold 3 would be in use but when full power is desired movement of the throttle to full open position would render the auxiliary system leading to manifold 9 operative also. Throttling is at all times controlled by adjustment of the governor I2. When 2| is closed and 23 opened, metering device being of the positive displacement type will not operate. This in turn will preclude shaft l8 from rotating and render metering device I6 inoperative, the entire flow then passing through by-pass line 22.

While the metering devices above described are of the gear type, it is apparent from the teachings of the invention that other functional equivalents will suggest themselves, and could be employed, to achieve the same ultimate results. It is contemplated that all metering devices which will distribute liquid to a plurality of discharge lines in a manner to efiect a constant ratio of mass flow between the respective discharge lines, are within the spirit and scope of the invention.

Fig. 2 illustrates another embodiment of a fuel system which will achieve substantially the same results as the one previously described. Primed figures represent parts corresponding to those of Fig. 1. In this embodiment the discharge line from governor l2 branches at l in the same -manner as the other embodiment, a portion of the fuel being delivered to the main fuel system 3 by line l9 and the remaining portion to auxiliary manifold 9 by line 20. This line is provided with a valve 2| to render the auxiliary supply inoperative, when its use is not desired, and an orifice 25 which will meter a constant pro- 4 portion of the discharge of governor l2 to th auxiliary system. The ratio of flows through I l and 20' may be adjusted for optimum exhaust temperature bythe choice ofa suitable orifice. The fixed orifice 25 may comprise an orifice as shown at 25" in Fig. 4. which is adjustable by means of needle'valve 21; this orifice would be so designed that its area could be adjusted by manual operation of the needle 21 during static calibration tests of the engine to maintain the proper fuel flow ratio to the main and auxiliary fuel injectors. The area so found is held constant during operation of the engine, or until further calibration tests are made at a later date.

In lieu of a fixed orifice as previously described, the orifice could be made adjustable by usin an adjustable valve 25' as shown in Fig. 3 in place of the fixed orifice 25 shown in Fig. 2,

or the fixed orifice 25" shown in Fig. 4. Thisvalve 25' could be placed under manual control of the pilot, who would make the adjustments in accordance with observed indications of the temperature at the exhaust pipe to produce and maintain a fuel fiow such that a desired exhaust pipe temperature would be maintained. The-use of the adjustable valve 25 under control of the pilot is for the purpose of placing the ratio of fuel fiows to the two combustion stages wthin the control of the pilot during flight. The advantage of this device lies in being able to precisely regulate the temperature of the gases from the auxiliary combustion chamber to a fixed valve. When the area of the orifice is maintained constant during a flight through various altitudes at a R. P. M. of the engine, calculations show that the exhaust gas temperature would var slightly. Adjustable valve 25 would allow the pilot to rectifythis, getting an indication of the exhaust gas temperature on hisinstrument board from a thermo-' couple or a similar device in the engine exhaust pipe, neither the thermocouple or other device being here il lustrated, as not a part of the invention.

As an added refinement, however, to obtain more precise temperature control the valve could be automatically controlled by a barometric device 26, as shown in Fig. 3, which would function to regulate the auxiliary flow in response to I changes of air pressure which is related to the resultin changes in the exhaust pipe temperatures due to changes in altitude at constant turbine speed.

The governor I2 maintaining constant turbine R. P. M. has nothing to do with orifice 25, valve 25, and barometric device 26. The governor operates independently of the proposed metering device and any of its parts.

The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

What I claim is:

l. A reaction motor having a, first stage combustion chamber providin gases for driving a turbine, and one or more additional stage combustion chambers for additional power generation, a fuel injector for each of said chambers, a governor controlled fuel metering means for supplying fuel under pressure at a controlled rate, a distribution system between said metering means and said injectors, including a proportional fuel divider means operated in response to said fuel pressure, a supply conduit from said metering means to said divided means, a fuel delivery line for delivering one portion of the fuel from said divider means to the first stage fuel injector, a second fuel delivery line for delivering the remainder of the fuel from said divider means to the remaining fuel injectors, a by-pass conduit from said supply conduit to said first delivery line, cut-off means in said by-pass conduit and in said second fuel delivery line, and control means for operating said cut-off means simultaneously so that when one is open the other is closed, whereby all the fuel may be selectively delivered either directly through said by-pass conduit into said first stage injector or proportionally through said divider means and said delivery lines into said first and additional stage injectors, irrespective of the rate of the total fuel supply.

2. A reaction motor as defined in claim 1 wherein said proportional fuel divider means comprises a pair of positive displacement gear pumps having a common drive shaft and a common inlet connected to said supply conduit, the outlet of one of said gear pumps being connected to said first fuel delivery line and the outlet of said other gear pump being connected to said second fuel delivery line, whereby when the cutoff means in said by-pass conduit is closed the fuel supply will be delivered proportionally to the first and second delivery line in accordance with the relative capacities of said gear pumps, and when the cut-off means in said by-pass conduit is open, the fuel divider will not operate, but all the fuel will be delivered to said first delivery line.

3. A reaction motor as defined in claim 2 wherein the gear pump connected to said first fuel delivery line has a smaller capacity than said other gear pump, in accordance with the relatively smaller fuel requirements of the primary combustion chamber, when the fuel is being delivered through said fuel divider.

4. A reaction motor as defined in claim 1 wherein said proportional fuel divider means comprises an open passage to said first fuel delivery line, and a restricted passage to said second delivery line precalibrated to deliver the desired proportions of the fuel supplied through said delivery lines.

5. A reaction motor as defined in claim 4 wherein said restricted passage comprises a fixed restriction.

6. A reaction motor as defined in claim 4 wherein said restricted passage comprises an adjustable restriction, with means for presetting the adjustment during calibration under static tests.

7. A reaction motor as defined in claim 4 wherein said restricted passage comprises a metering valve for manual control during motor operation.

EDWARD M. BEDDING.

REFERENCES CITED The following references are of record in the file of this patent:

UNITED STATES PATENTS Number Name Date 2,238,905 Lysholm Apr. 22, 1941 2,409,176 Allen -Oct. 15, 1946

US640293A 1946-01-10 1946-01-10 Fuel control system for turbojet engines Expired - Lifetime US2566373A (en)

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Cited By (41)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2640316A (en) * 1949-11-07 1953-06-02 Westinghouse Electric Corp Ignition apparatus for turbojet afterburners
US2656676A (en) * 1949-02-19 1953-10-27 Mcdonnell Aircraft Corp Ram jet engine
US2659196A (en) * 1949-08-09 1953-11-17 United Aircraft Corp Centrifugal fuel supply means for jet engine afterburners
US2671186A (en) * 1952-02-12 1954-03-02 K L G Sparking Plugs Ltd Electric spark discharge device
US2674843A (en) * 1948-01-28 1954-04-13 Rolls Royce Temperature controlled fuel system of gas-turbine engines having reheat combustion equipment
US2679726A (en) * 1952-03-27 1954-06-01 United Aircraft Corp Gas turbine power plant
US2683349A (en) * 1949-11-17 1954-07-13 Lucas Ltd Joseph Jet engine combustion system having burner in the jet pipe and controlling means therefor
US2693081A (en) * 1951-04-04 1954-11-02 Allen S Atkinson Apparatus for controlling gas turbine engines during transient operation
US2699646A (en) * 1949-06-30 1955-01-18 Gen Electric Gas turbine power plant having variable exhaust area and control system therefor
US2700275A (en) * 1948-12-21 1955-01-25 Niles Bement Pond Co Fuel control apparatus for turbojet engines
US2705868A (en) * 1948-02-19 1955-04-12 Power Jets Res & Dev Ltd Combustion apparatus
US2706888A (en) * 1949-03-10 1955-04-26 Rolls Royce Pump arrangements for gas-turbine engine fuel systems
US2707372A (en) * 1947-06-11 1955-05-03 Lockheed Aircraft Corp Afterburner apparatus for turbo jet engines having movable flame holder means
US2742755A (en) * 1949-11-14 1956-04-24 Rolls Royce Fuel system for pilot burners of gasturbine engines reheat equipment
US2744384A (en) * 1952-08-09 1956-05-08 United Aircraft Corp Burner construction for high velocity gases
US2745249A (en) * 1946-10-22 1956-05-15 Ryan Aeronautical Co Reheater and fuel vaporizer for jet propulsion engines
US2753686A (en) * 1951-05-16 1956-07-10 United Aircraft Corp Ramjet fuel regulator
US2754655A (en) * 1950-07-13 1956-07-17 Kellogg M W Co Thrust cylinder with integrated turbine
US2774215A (en) * 1949-04-22 1956-12-18 Bendix Aviat Corp Tailpipe or afterburning control for turbojet engines
US2778191A (en) * 1948-06-03 1957-01-22 Bendix Aviat Corp Tail pipe or afterburning control for turbojet engines
US2780915A (en) * 1951-12-05 1957-02-12 Solar Aircraft Co Fuel distribution system for jet engine and afterburner
US2818703A (en) * 1954-07-01 1958-01-07 Gen Electric Jet engine fuel, pressure ratio, and nozzle area control
US2821065A (en) * 1952-07-05 1958-01-28 Gen Electric Exhaust temperature regulator for gas turbine power-plant
US2823520A (en) * 1951-05-10 1958-02-18 Spalding Dudley Brian Combustion equipment and gas turbine plant
US2828606A (en) * 1950-11-18 1958-04-01 United Aircraft Corp Afterburner fuel metering device for turbojet engines
US2835108A (en) * 1951-07-17 1958-05-20 Solar Aircraft Co Variable area flametholder for afterburner
US2857739A (en) * 1951-04-06 1958-10-28 Pratt & Whitney Co Inc Control system for turbo-jet engine
US2865166A (en) * 1952-06-11 1958-12-23 Gen Motors Corp Afterburner fuel system with pump unloader
US2882679A (en) * 1950-12-22 1959-04-21 Gen Motors Corp Augmenter type afterburner for jet propelled aircraft
US2896408A (en) * 1953-09-23 1959-07-28 Republic Aviat Corp Turbojet convertible to a ramjet
US2932946A (en) * 1952-05-20 1960-04-19 Rolls Royce Fuel system for gas turbine engine including hydraulically driven auxillary pump
US2953899A (en) * 1951-05-17 1960-09-27 Honeywell Regulator Co Fuel flow controller for gas turbines and jet propulsion units
US2960155A (en) * 1953-05-26 1960-11-15 Bendix Corp Afterburner fuel metering control
US2965065A (en) * 1955-06-15 1960-12-20 Walter H Tinker Hydraulic jet propulsion units for boats
US2988875A (en) * 1956-11-21 1961-06-20 United Aircraft Corp Afterburner fuel control having multiple sets of nozzles
US3029599A (en) * 1953-01-21 1962-04-17 Chandler Evans Corp Jet engine afterburner fuel control
US3067576A (en) * 1959-08-28 1962-12-11 Gen Motors Corp Co-ordinated main and afterburner fuel controls for a turbojet
US3174281A (en) * 1962-12-14 1965-03-23 Gen Motors Corp Afterburner fuel control
US3232053A (en) * 1963-09-13 1966-02-01 Bendix Corp Fuel feed and power control apparatus for combustion engines
US3379009A (en) * 1964-06-06 1968-04-23 Bristol Siddeley Engines Ltd Combustion system for fluid fuel
RU2531110C2 (en) * 2010-06-29 2014-10-20 Дженерал Электрик Компани Gas-turbine unit and unit with injector vanes (versions)

Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2238905A (en) * 1937-05-14 1941-04-22 Milo Ab Gas turbine plant
US2409176A (en) * 1942-04-10 1946-10-15 Allis Chalmers Mfg Co Gas turbine system

Patent Citations (2)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2238905A (en) * 1937-05-14 1941-04-22 Milo Ab Gas turbine plant
US2409176A (en) * 1942-04-10 1946-10-15 Allis Chalmers Mfg Co Gas turbine system

Cited By (42)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2745249A (en) * 1946-10-22 1956-05-15 Ryan Aeronautical Co Reheater and fuel vaporizer for jet propulsion engines
US2707372A (en) * 1947-06-11 1955-05-03 Lockheed Aircraft Corp Afterburner apparatus for turbo jet engines having movable flame holder means
US2674843A (en) * 1948-01-28 1954-04-13 Rolls Royce Temperature controlled fuel system of gas-turbine engines having reheat combustion equipment
US2705869A (en) * 1948-02-19 1955-04-12 Power Jets Res & Dev Ltd Combustion apparatus
US2705868A (en) * 1948-02-19 1955-04-12 Power Jets Res & Dev Ltd Combustion apparatus
US2778191A (en) * 1948-06-03 1957-01-22 Bendix Aviat Corp Tail pipe or afterburning control for turbojet engines
US2700275A (en) * 1948-12-21 1955-01-25 Niles Bement Pond Co Fuel control apparatus for turbojet engines
US2656676A (en) * 1949-02-19 1953-10-27 Mcdonnell Aircraft Corp Ram jet engine
US2706888A (en) * 1949-03-10 1955-04-26 Rolls Royce Pump arrangements for gas-turbine engine fuel systems
US2774215A (en) * 1949-04-22 1956-12-18 Bendix Aviat Corp Tailpipe or afterburning control for turbojet engines
US2699646A (en) * 1949-06-30 1955-01-18 Gen Electric Gas turbine power plant having variable exhaust area and control system therefor
US2659196A (en) * 1949-08-09 1953-11-17 United Aircraft Corp Centrifugal fuel supply means for jet engine afterburners
US2640316A (en) * 1949-11-07 1953-06-02 Westinghouse Electric Corp Ignition apparatus for turbojet afterburners
US2742755A (en) * 1949-11-14 1956-04-24 Rolls Royce Fuel system for pilot burners of gasturbine engines reheat equipment
US2683349A (en) * 1949-11-17 1954-07-13 Lucas Ltd Joseph Jet engine combustion system having burner in the jet pipe and controlling means therefor
US2754655A (en) * 1950-07-13 1956-07-17 Kellogg M W Co Thrust cylinder with integrated turbine
US2828606A (en) * 1950-11-18 1958-04-01 United Aircraft Corp Afterburner fuel metering device for turbojet engines
US2882679A (en) * 1950-12-22 1959-04-21 Gen Motors Corp Augmenter type afterburner for jet propelled aircraft
US2693081A (en) * 1951-04-04 1954-11-02 Allen S Atkinson Apparatus for controlling gas turbine engines during transient operation
US2857739A (en) * 1951-04-06 1958-10-28 Pratt & Whitney Co Inc Control system for turbo-jet engine
US2823520A (en) * 1951-05-10 1958-02-18 Spalding Dudley Brian Combustion equipment and gas turbine plant
US2753686A (en) * 1951-05-16 1956-07-10 United Aircraft Corp Ramjet fuel regulator
US2953899A (en) * 1951-05-17 1960-09-27 Honeywell Regulator Co Fuel flow controller for gas turbines and jet propulsion units
US2835108A (en) * 1951-07-17 1958-05-20 Solar Aircraft Co Variable area flametholder for afterburner
US2780915A (en) * 1951-12-05 1957-02-12 Solar Aircraft Co Fuel distribution system for jet engine and afterburner
US2671186A (en) * 1952-02-12 1954-03-02 K L G Sparking Plugs Ltd Electric spark discharge device
US2679726A (en) * 1952-03-27 1954-06-01 United Aircraft Corp Gas turbine power plant
US2932946A (en) * 1952-05-20 1960-04-19 Rolls Royce Fuel system for gas turbine engine including hydraulically driven auxillary pump
US2865166A (en) * 1952-06-11 1958-12-23 Gen Motors Corp Afterburner fuel system with pump unloader
US2821065A (en) * 1952-07-05 1958-01-28 Gen Electric Exhaust temperature regulator for gas turbine power-plant
US2744384A (en) * 1952-08-09 1956-05-08 United Aircraft Corp Burner construction for high velocity gases
US3029599A (en) * 1953-01-21 1962-04-17 Chandler Evans Corp Jet engine afterburner fuel control
US2960155A (en) * 1953-05-26 1960-11-15 Bendix Corp Afterburner fuel metering control
US2896408A (en) * 1953-09-23 1959-07-28 Republic Aviat Corp Turbojet convertible to a ramjet
US2818703A (en) * 1954-07-01 1958-01-07 Gen Electric Jet engine fuel, pressure ratio, and nozzle area control
US2965065A (en) * 1955-06-15 1960-12-20 Walter H Tinker Hydraulic jet propulsion units for boats
US2988875A (en) * 1956-11-21 1961-06-20 United Aircraft Corp Afterburner fuel control having multiple sets of nozzles
US3067576A (en) * 1959-08-28 1962-12-11 Gen Motors Corp Co-ordinated main and afterburner fuel controls for a turbojet
US3174281A (en) * 1962-12-14 1965-03-23 Gen Motors Corp Afterburner fuel control
US3232053A (en) * 1963-09-13 1966-02-01 Bendix Corp Fuel feed and power control apparatus for combustion engines
US3379009A (en) * 1964-06-06 1968-04-23 Bristol Siddeley Engines Ltd Combustion system for fluid fuel
RU2531110C2 (en) * 2010-06-29 2014-10-20 Дженерал Электрик Компани Gas-turbine unit and unit with injector vanes (versions)

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