US2683349A - Jet engine combustion system having burner in the jet pipe and controlling means therefor - Google Patents

Jet engine combustion system having burner in the jet pipe and controlling means therefor Download PDF

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US2683349A
US2683349A US192411A US19241150A US2683349A US 2683349 A US2683349 A US 2683349A US 192411 A US192411 A US 192411A US 19241150 A US19241150 A US 19241150A US 2683349 A US2683349 A US 2683349A
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burner
jet
control
controlling
jet pipe
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US192411A
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Lawrence Owen Napier
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Joseph Lucas Ltd
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Joseph Lucas Ltd
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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

y 1954 o. N. LAWRENCE 83,349
JET ENGINE COMBUSTION SYSTEM HAVING BURNER IN THE JET PIPE AND CONTROLLING MEANS THEREFOR Filed Oct, 27, 1950 W zZl/UelbtOl Fig.2
Patented July 13, 1954 TENT OFFICE JET ENGINE COMBUSTION SYSTEM HAVING BURNER IN THE JET PIPE AND CONTROL- LING MEANS THEREFOR Owen Napier Lawrence, Dorridge, England, as-
signor to Joseph Lucas Limited, Birmingham,
England Application October 27, 1950, Serial No. 192,411
Claims priority, application Great Britain November 17, 1949 3 Claims.
This invention relates to the combustion systems of jet engines, of the kind in which the system (or each of an associated group of such systems) comprises an elongated combustion chamber which at one end is provided with a liquid fuel burner and is adapted for the admission thereto of air from a blower, the other end terminating in an extension which forms the jet pipe, and in which there is mounted, between the adjacent ends of the combustion chamber and jet pipe, the rotor of a turbine which actuates the blower.
In some such systems provision is made for admitting additional liquid fuel to a. burner in the jet pipe, when it is required to augment the energy of the gases therein after passing the turbine and before discharge to the atmosphere. It is undesirable however, to admit the additional fuel at an excessive rate. The permissible maximum rate of supply is not a fixed quantity, but is dependent on contemporary working conditions in the system, and the object of the present invention is to enable the rate of supply to be automatically controlled in a satisfactory manner.
The invention comprises the combination with the system, of means for controlling the additional fuel under the joint action of the pressure of the air at the blower inlet, and the pressure of the gas in the jet pipe between the turbine and the point of admission of the additional fuel.
The invention also comprises means as specified in the preceding paragraph, in which the relative effects of the said pressures is variable unde manual control.
In the accompanying drawing, Figures 1 and 2 are diagrams illustrating two embodiments of the invention.
Referring to Figure 1, the principal combustion chamber of a jet-engine is represented by a. In this is mounted a liquid fuel burner b which is supplied by a pump at a variable rate in response to the control of the pilot and. of conditions associated with the working of th engine, such as speed, 01' atmospheric pressure. The blower for air required in the combustion chamber is represented by c and this is driven by a turbine 11 situated at the discharge end of the combustion chamber. In the extension or jetpipe e of the combustion chamber is arranged a burner f by which the additional fuel is supplied, and it is the control of this fuel which forms the subject of the present invention.
In the example illustrated 'at Figure l, the extra fuel is supplied by a rotary pump 5/ of the swash plate type, the output of the pump being variable by adjustment of the obliquity of the swash plate h under the action of a fuel-operated servo-mechanism. This mechanism is of known form and comprises a cylindrical chamber 2' containing a piston 7' loaded by a spring k and connected to the swash plate by a rod m. One end of the cylinder 2' is in communication with the discharge passage n of the pump by way of a passage 0, and the other end is in communication with the passage 11 by way of a restricted passage 21. The end of the cylinder 11 containing the spring communicates with the seating q of a control valve 1" by way of a passage (1;. The arrangement is such that when the valve 1" is closed, the liquid pressures acting on opposite sides of the piston are balanced, and the spring is then moves the piston in the direction for increasing the pump output. When the valve 7 is opened, a preponderating liquid pressure acts on the side of the piston in opposition to the spring and moves the piston in the direction for reducing the pump output.
The valve 1' is mounted on a lever s which is carried by a flexible diaphragm t, or is pivotally mounted on a partition, which separates two compartments in, 'u, in a chambered body part, the compartment it being in communication with the pump inlet, or a sump. The end of the lever which extends into the compartment 1) is connected by a link 2 to a diaphragm 3. One side of the diaphragm is subject to the pressure of the air at the inlet side of the blower c, a passage 4 being provided between the blower inlet and the compartment 12. The other side of the diaphragm is subject to the pressure of the gases at the exit side of the turbine at, a passage 5 being provided for effecting communication between the said side of the turbine and diaphragm.
In the chamber 1) is also contained another lever 6 loaded by a spring 1 and evacuated capsule 8. Between the two levers s, 5 is arranged a slidable abutment 9 by which force can be transmitted from one lever to the other, and the position of the abutment relatively to the lever is movable under the manual control of the pilot through a lever iii. By appropriately adjusting the position of the abutment 9 the forces acting on the valve lever 1' can be varied as desired.
Assuming that the pressures acting on the diaphragm 3 and capsule 8 are such that the valve 1' is slightly open, and that the piston j is in equilibrium at a position intermediate the ends of the chamber 2', it will be seen that the pressure exerted on the left hand side of the piston by liquid admitted to the corresponding end of the chamber through the passage 0 is balanced by the pressure exerted on the other side of the piston by the combined effect of the spring k: and the liquid admitted to the right hand end of the chamber through the restricted passage p, the liquid pressure in the last mentioned end of the chamber being partially relieved by leakage of liquid past the valve r. If now the valve r is opened to a greater extent by relative variation of the pressures acting on the diaphragm it and capsule 3, the liquid pressure in the right hand end of the chamber 2' will be further relieved, and the piston will be moved against the action of the spring is by the preponderating liquid pressure in the left hand end of the chamber to cause reduction of the pump output until a new condition of equilibrium is reached. Closing movement of the valve 1' by relative variation of the pressures acting on the diaphragm 3 and capsule 8 has a reverse effect to that above described, and enables the spring is to move the piston 7 for increasing the pump output. Movement of the abutment a under the control of the pilot varies the efiect of the capsule 8 on the lever s without varying the effect of the diaphragm 3 on this lever.
It will be apparent therefore, that by the arrangement above described, the rate of supply of fuel from the pump g to the burner f is automatically controllable by the pressures acting on the diaphragm 3, the air pressure acting on the capsule 8 and the load exerted by the spring 7, the effect on the valve 1" being variable under the control of the pilot through the lever it and abutment 9.
The embodiment shown in Figure 2 differs from that shown in Figure 1, in that the one pump g supplies liquid fuel to both of the burners b, f. In this example the servo-mechanism i, 7, 7c, 0, go above described serves mainly to control the supply of fuel to the burner 19, and this is effected in response to any control device positioned as indicated by 13 in the diagram between the servomechanism and the burner b. The supply of fuel from the pump to the burner f is mainly controlled by a throttle H, which is actuated by a second servo-mechanism indicated generically by i2 and which. is essentially similar to the one above described with reference to Figure 1. The servo-mechanism 22 is controlled by a similar valve mechanism to that already described, similar reference characters being used in Figure 2 to identify the. valve mechanism.
Having thus described my invention what I claim as new and desire to secure by Letters Patent is:
1. A jet-engine combustion system having in combination anelongated combustion chamber, a main burner in the combustion chamber, a blower for supplying air under pressure to one end of the combustion chamber, a turbine at the other end of the combustion chamber, a jet pipe extending from the last mentioned end of the combustion chamber, an auxiliary burner in the jet pipe, means for supplying liquid fuel to the auxiliary burner, a liquid-operated servo-mechanism for determining the quantity of liquid fuel supplied to the auxiliary burner, control means responsive to'theairpressure at the blower inlet,
additional control means responsive to the gas pressure in the jet pipe between the turbine and the auxiliary burner, and a control member responsive-to the combined action of the two control means for controlling the action of said servomechanism.
2. A jet-engine cmbustion system as claimed in claim 1 and having manually adjustable means for varying the combined effect of said two control means on said control member, said manually adjustable means being formed in part by a motion-transmitting abutment through which one of said two control means acts on said control member.
3. A jet-engine combustion system as claimed in claim 1, in which the means for supplying liquid fuel to said auxiliary burner comprises a pump which serves also for supplying fuel to main burner, and in which a throttle operable by the liquid-operated servo-mechanism is provided for controlling the quantity of liquid fuel supplied to said auxiliary burner.
References Cited in the file of this patent UNITED STATES PATENTS Number Name Date 2,438,663 Greenland Mar. 30, 1948 2,498,939 Bobier, Jr. Feb. 23, 1950 2,503,048 Ifield Apr. 4, 1950 2,506,611 Neal et al May 9, 1950 2,537,772 Lundquist et a1. Jan. 9, 1951 2,545,703 Orr, Jr Mar. 20, 1951 2,566,373 Bedding Sept. 4, 1951 2,580,962 Sdille Jan. 1, i952 FOREIGN PATENTS Number Country Date 578,311 Great Britain June 1946
US192411A 1949-11-17 1950-10-27 Jet engine combustion system having burner in the jet pipe and controlling means therefor Expired - Lifetime US2683349A (en)

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Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2818703A (en) * 1954-07-01 1958-01-07 Gen Electric Jet engine fuel, pressure ratio, and nozzle area control
US2820435A (en) * 1950-11-18 1958-01-21 United Aircraft Corp Pressure responsive control device
US2844936A (en) * 1955-07-15 1958-07-29 Ca Nat Research Council Control of reheat in turbojet engines
US2872781A (en) * 1953-05-05 1959-02-10 Rolls Royce Gas-turbine engine reheat fuel supply system with air turbine driven fuel pump
US2872784A (en) * 1953-03-31 1959-02-10 Lucas Industries Ltd Liquid fuel control means for jet engines
DE1055299B (en) * 1955-07-08 1959-04-16 Bendix Aviat Corp Afterburner control device for gas turbine jet engines
US2933887A (en) * 1953-07-16 1960-04-26 Rolls Royce Compound gas turbine engine with control for low-pressure rotor
US2988875A (en) * 1956-11-21 1961-06-20 United Aircraft Corp Afterburner fuel control having multiple sets of nozzles
US2989850A (en) * 1956-02-23 1961-06-27 Bendix Corp Gas turbine fuel control system for preventing compressor stall
US3094072A (en) * 1957-12-09 1963-06-18 Arthur R Parilla Aircraft, missiles, missile weapons systems, and space ships
US3234730A (en) * 1959-04-21 1966-02-15 Bendix Corp Dual afterburner manifold proportioning control
US3572038A (en) * 1968-07-29 1971-03-23 Bendix Corp Thrust control mechanism

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB578311A (en) * 1943-06-22 1946-06-24 Alan Arnold Griffith Improvements in or relating to aircraft
US2438663A (en) * 1945-02-12 1948-03-30 Hobson Ltd H M Fuel injection system for internalcombustion engines
US2498939A (en) * 1948-11-01 1950-02-28 George M Holley Gas turbine tail burner fuel control
US2503048A (en) * 1945-12-27 1950-04-04 Lucas Ltd Joseph Means for controlling the flow of liquid fuel to prime movers
US2506611A (en) * 1948-03-02 1950-05-09 Westinghouse Electric Corp Fuel control for aviation gas turbine power plants
US2537772A (en) * 1944-11-30 1951-01-09 Wright Aeronautical Corp Turbo-prop exhaust nozzle control system utilizing impact and exhaust pressures as parameters
US2545703A (en) * 1947-03-17 1951-03-20 George M Holley Gas turbine temperature control responsive to air and fuel flow, compressor intake and discharge temperature and speed
US2566373A (en) * 1946-01-10 1951-09-04 Edward M Redding Fuel control system for turbojet engines
US2580962A (en) * 1945-06-13 1952-01-01 Rateau Soc Control means for the nozzle outlet area of jet engines

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB578311A (en) * 1943-06-22 1946-06-24 Alan Arnold Griffith Improvements in or relating to aircraft
US2537772A (en) * 1944-11-30 1951-01-09 Wright Aeronautical Corp Turbo-prop exhaust nozzle control system utilizing impact and exhaust pressures as parameters
US2438663A (en) * 1945-02-12 1948-03-30 Hobson Ltd H M Fuel injection system for internalcombustion engines
US2580962A (en) * 1945-06-13 1952-01-01 Rateau Soc Control means for the nozzle outlet area of jet engines
US2503048A (en) * 1945-12-27 1950-04-04 Lucas Ltd Joseph Means for controlling the flow of liquid fuel to prime movers
US2566373A (en) * 1946-01-10 1951-09-04 Edward M Redding Fuel control system for turbojet engines
US2545703A (en) * 1947-03-17 1951-03-20 George M Holley Gas turbine temperature control responsive to air and fuel flow, compressor intake and discharge temperature and speed
US2506611A (en) * 1948-03-02 1950-05-09 Westinghouse Electric Corp Fuel control for aviation gas turbine power plants
US2498939A (en) * 1948-11-01 1950-02-28 George M Holley Gas turbine tail burner fuel control

Cited By (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2820435A (en) * 1950-11-18 1958-01-21 United Aircraft Corp Pressure responsive control device
US2872784A (en) * 1953-03-31 1959-02-10 Lucas Industries Ltd Liquid fuel control means for jet engines
US2872781A (en) * 1953-05-05 1959-02-10 Rolls Royce Gas-turbine engine reheat fuel supply system with air turbine driven fuel pump
US2933887A (en) * 1953-07-16 1960-04-26 Rolls Royce Compound gas turbine engine with control for low-pressure rotor
US2818703A (en) * 1954-07-01 1958-01-07 Gen Electric Jet engine fuel, pressure ratio, and nozzle area control
DE1055299B (en) * 1955-07-08 1959-04-16 Bendix Aviat Corp Afterburner control device for gas turbine jet engines
US2844936A (en) * 1955-07-15 1958-07-29 Ca Nat Research Council Control of reheat in turbojet engines
US2989850A (en) * 1956-02-23 1961-06-27 Bendix Corp Gas turbine fuel control system for preventing compressor stall
US2988875A (en) * 1956-11-21 1961-06-20 United Aircraft Corp Afterburner fuel control having multiple sets of nozzles
US3094072A (en) * 1957-12-09 1963-06-18 Arthur R Parilla Aircraft, missiles, missile weapons systems, and space ships
US3234730A (en) * 1959-04-21 1966-02-15 Bendix Corp Dual afterburner manifold proportioning control
US3572038A (en) * 1968-07-29 1971-03-23 Bendix Corp Thrust control mechanism

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