US3124933A - Leroy stram - Google Patents

Leroy stram Download PDF

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US3124933A
US3124933A US3124933DA US3124933A US 3124933 A US3124933 A US 3124933A US 3124933D A US3124933D A US 3124933DA US 3124933 A US3124933 A US 3124933A
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fuel
combustion chamber
engine
starting
flame
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02CGAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/26Starting; Ignition
    • F02C7/264Ignition

Description

March 17, 1964 LE ROY STRAM, JR, ETAL 3,124,933
STARTING SYSTEM Filed March 9, 1959 I5 Sheets-Sheet 1 I pan/ 28 l5 4 I I2 26 a,
INVENTORS HENRY C. MASKEY Le ROY STRAM.JR. BY SEYMOUR A. GENDEN March 17, 1964 LE ROY STRAM, JR., ETAL 3,124,933
I STARTING SYSTEM Filed March 9, 1959 3 Sheets-Sheet 2 55 2i, |e lo FIG.4. T1
57 & J 25 60 GI 27 INVENTORS HENRY C. MASKEY Le ROY STRAM, JR. SEYMOUR A.G ENDEN March 17, 1964 LE ROY STRAM, JR., ETAL 3,
STARTING SYSTEM Filed March 9, 1959 3 Sheets-Sheet 5 F G 5 Y OXYGEN SUPPLY 3? OXYGEN sou-mom OXYGEN ORFICE oxvssu 39- 39 CHECK VALVE 35 35 I FUEL rowan 24vo.c. 3 v VALVE FUEL onncs 4 FUEL SOLENOID 40 FUEL PUMP To MAIN FUEL SOURCE 65 MAIN FUEL 55-" SUPPLY BOOST PUMP 56 I REVERSE FLOW CH ECK VALVE PROPYL NITRATE MAIN FUEL 68 SHUT- OFF VALVE TO ENGINE RELA BOX 9 1 THROTTLE o CONTROL 6 ACTUATOR MOTOR T cousous Q 4vo.c. POWER.SOURcE E FIG 6 INVENTORS HENRY C. MASKEY L8 ROY STRAM,JR.
BY SEYMOUR A .GENDEN T ATTQRQEYS United States Patent O 3,124,933 STARTING SYSTEM Le Roy Strain, .In, Royal Oak, Seymour A. Genden, Oak Park, and Henry C. Maskey, Grosse Pointe Farms, Mich assignors to Qontinental Aviation and Engineering Corporation, Detroit, Mich, a corporation of Virginia Filed Mar. 9, 1959, Ser. No. 798,218 2 Claims. (Cl. 6039.82)
Our invention relates to high altitude starting systems for turbojet engines and more particularly to a system for producing a high energy flame within a turbojet combustion chamber to ignite engine fuel for high altitude air starting.
Originally, small turbojet engines for which the present system was devised, could only be started by using high vapor pressure fuels, such as aviation gasoline, and at ambient conditions approximating sea level static operation. Once the engine was over its idling speed, the fuel supply could be switched to the l P type of fuels, and the engine would operate satisfactorily over a wide range of ambient temperatures, inlet ram pressure ratios and altitudes.
However, installation of these engines in target drones and the like required an ignition system capable of starting the engine at certain minimum altitudes and higher true air speeds. Various ignition systems have been in vestigated involving use of high energy electric coils and primer fuel sprayed into the air induction system, the primer fuel being ignited and carried into the combustor primary combustion zone where fuel spraying from the fuel slinger became ignited. These systems only proved satisfactory up to about 25,000 feet altitude, and 300 mph true air speed at a temperature of about 45 F. At higher altitudes and speeds and lower ambient temperatures, combustion instability during starting grows more severe and the energy required for ignition proves to be far too great for practical uses of such systems, and even the most advantageous configurations of igniter coils, igniter plugs, and primary fuel spray nozzles could not solve the problems, particularly when it became desirable to effect air starting at altitudes from 40,000 to 75,000 feet at speeds of about Mach .5 to .9.
Another difiiculty encountered in starting these small turbojet engines when used in air-dropped target drones is that the starting procedure calls for an experienced operator in the launch aircraft, who must properly accelerate the engine by means of some type of throttle actuator While maintaining a maximum exhaust temperature.
Basically, an object of the present invention is to solve high altitude and high speed starting problems by providing sustained high energies within the combustion chamber for igniting the fuel delivered thereto.
Another object of the invention is to make possible turbojet engine starting at high altitudes and high true air speeds by devising an ignition system capable of introducing an initial high energy flame within the combustion chamber for igniting engine fuel.
A further object of the invention is to provide a reliable high altitude and high speed starting system for turbojet engines by constructing a combustion chamber having means for producing a sustained high energy flame during the critical starting and initial acceleration period, said high energy flame being independent of the primary fuel and induction air conditions.
For a more complete understanding of the invention, reference may be had to the accompanying drawings illustrating preferred embodiments of the invention in which like reference characters refer to like parts throughout the several views and in which 3,124,933 Patented Mar. 17, 1964 FIG. 1 is a partially sectioned illustration of a preferred turbojet engine embodying the present invention.
FIG. 2 is a partially sectioned illustration of a preferred turbojet engine embodying another modification of the invention.
FIG. 3 is an enlarged cross-sectioned view of a portion of the engine of FIG. 1.
FIG. 4 is an enlarged cross-sectioned view of a portion of the engine of FIG. 2.
FIG. 5 is a diagrammatic illustration of the starting system of the engine of FIG. 1, and
FIG. 6 is a diagrammatic illustration of the starting system of the engine of FIG. 2.
Referring to FIGS. 1 and 3, a preferred type of small turbojet engine 10 is illustrated as having a turbine 11 drivingly connected with a single stage compressor 12 through a drive shaft 13, an annular combustion chamber 14 defined generally by an inner combustor shell 15 and an outer combustor shell 16, an annular compressed air chamber 17 defined generally by the outer combustor shell 16 and an engine housing structure 13, the engine housing structure 18 including an air intake structure 19 and an exhaust nozzle structure 20.
Compressed air is delivered by the compressor 12 as indicated by the arrows through an annular intake passage 25 to the compressed air chamber 17 and thence radially inward and rearward through a swirl vane structure 26 into what is termed the primary combustion zone; axially rearward, radially inward and forward to inside the inner combustor shell 15 and through ports 27 into the rear side of the primary combustion zone; and radially inward through ports 28 in the outer combustor shell 16 into what is termed the secondary combustion zone.
A fuel slinger 29 is carried in the drive shaft 13 and operates by centrifugal force to spray fuel radially outwardly into the primary combustion zone, the fuel being conducted to the slinger 29 through a conduit 30 from the engine supply system (not shown).
The starting system for the engine 10 generally is illustrated in FIG. 5 as preferably comprising a pair of torch igniters 35, an oxygen supply 36 from which oxygen is delivered through a solenoid shut-off valve 37, orifices 38, and check valves 39 to the torches 35. Fuel from the engine main fuel source is delivered by a fuel pump 40 through a solenoid shut-off valve .41, orifices 42, and check valves 43 to the torches 35. The orifices 39 and 43 are included in the system to respectively govern the flow characteristics of the oxygen and fuel.
One of the torches 35 is illustrated in FIG. 3 as comprising a tube 45 mounted on the engine housing 18 and having a nozzle 46 operable to direct a flame directly into the combustion chamber 14 through a hole 47 in the outer combustor shell 16. The oxygen and fuel are respectively conducted into the tube 45 through any suitable inlet connections 48 and 49 from their respective sources, and the resulting combustible mixture is ignited by any suitable means such as a spark igniter 50. It will thus be apparent that the igniter operates independently of conditions in either the air induction system or the engine fuel delivery system.
The high energy flame produced by the torch 35 is directed into the primary combustion zone of the combustion chamber where it will ignite the engine fuel-air mixture and sustain combustion in the combustion chamber for as long as desired, being limited only by the supply of oxygen carried. Combustion can readily be sustained throughout the critical acceleration period of the engine until conditions are reached under which the normal engine fuel and air supply systems will maintain combustion. Depending on the supply of oxygen,
this system will also operate to provide for multiple starts if desired.
A modification of the invention is illustrated in FIGS. 2, 4 and 6, in which the same engine is involved, but instead of the torches 35, pyrotechnic flares 55 are utilized for starting. One such flare 55 is illustrated in FIG. 4 as comprising a metal sleeve 56 mounted on the engine housing 18, a combustible plastic tube 57 carried in the sleeve 56 and projecting therefrom through a hole 58 in the outer combustion shell 16 into the primary combustion zone of the combustion chamber 14. The tube 57 is packed with a charge 59 of a suitable composite type rocket propellant including its own oxy dizer, such as polysulfide-ammonium nitrate. The tube 57 has an open end 60 and preferably carries a glow wire 61 in contact with the rocket propellant charge 59 and suitably connected by means of an electrical plug connection 62 to any suitable electrical system operable to heat the glow wire 61 for initially igniting the propellant charge 59.
FIG. 6 illustrates a preferred fuel and electrical system for use with the pyrotechnic flare. Between a main engine fuel supply 65 and the engine is a measured supply 66 of a suitable primer fuel such as propyl nitrate, the primer fuel being delivered to the engine first under the pressure of the main fuel delivered through a fuel shut-01f valve 67 by a boost pump 68. The flares 55, the fuel shut-off valve 67, the boost pump 68 and a throttle actuator motor 69 are operated through a common electrical system, the throttle actuator motor 69 operating to regulate fuel flow to prevent quenching of the flare flame during the critical acceleration period.
It will be seen that the pyrotechnic flare method of starting will be operable as a one-start system only, and the time of high energy flame supply will be dependent on the size and type of propellant charge 59. The propyl nitrate primer fuel has been found to be advantageous for starting since it has a much faster flame propagation rate than that of the hydrocarbon fuels, the low pressures and temperatures experienced at high altitudes materially reducing the flame propagation rates of the hydrocarbon fuels. The propylnitrate thus prevents afterburning and combustion instability during starting.
Although we have described only two embodiments of the invention, it will be apparent to one skilled in the art to which the invention pertains that various changes and modifications may be made therein without departing from the spirit of the invention or the scope of the appended claims.
We claim:
1. A high altitude starting system for a turbojet engine and comprising (a) a housing structure,
(b) a combustor structure disposed within said housing structure and defining a combustion chamber,
(0) a fuel supply means delivering a primary fuel stream to said combustion chamber,
(0') air induction means delivering air to said combustion chamber to provide high energy fuel, and
(e) a flame producer operable to provide a high energy flame within said combustion chamber for initiating combustion of said primary fuel,
(f) said flame producer comprising a pyrotechnic flare supported within said combustion chamber and having a charge of high energy solid rocket fuel including an oxidizer and means igniting said rocket fuel,
(g) said flare comprising a metal sleeve secured to said housing structure and having an open end disposed adjacent said combustor structure and exteriorly of said combustion chamber,
(It) said flare further comprising a combustible tube carrying said charge of rocket fuel and carried within said sleeve and extending from said open end through said combustor structure and into said combustion chamber, whereby said sleeve is disposed substantially entirely exteriorly of said combustion chamber and said combustible tube is disposed substantially entirely inside said combustion chamber.
2. The starting system as in claim 1 and in which said igniting means comprises a glow wire carried by said tube and adjacent said charge of rocket fuel, and an electrical ignition system including means heating said fglow wire to initiate combustion of said charge of rocket uel.
References Cited in the file of this patent UNITED STATES PATENTS 2,518,882 Goddard Aug. 15, 1950 2,561,670 Miller July 24, 1951 2,595,759 Buclkland May 6, 1952 2,760,340 Seglem Aug. 28, 1956 2,784,553 De Croso Mar. 12, 1957 2,856,755 Szydlowski Oct. 21, 1958 2,937,501 Trousse May 24, 1960 FOREIGN PATENTS 627,722 Great Britain Aug. 15, 1949 714,800 Great Britain Sept. 1, 1954

Claims (1)

1. A HIGH ALTITUDE STARTING SYSTEM FOR A TURBOJET ENGINE AND COMPRISING (A) A HOUSING STRUCTURE, (B) A COMBUSTOR STRUCTURE DISPOSED WITHIN SAID HOUSING STRUCTURE AND DEFINING A COMBUSTION CHAMBER, (C) A FUEL SUPPLY MEANS DELIVERING A PRIMARY FUEL STREAM TO SAID COMBUSTION CHAMBER, (D) AIR INDUCTION MEANS DELIVERING AIR TO SAID COMBUSTION CHAMBER TO PROVIDE HIGH ENERGY FUEL, AND (E) A FLAME PRODUCER OPERABLE TO PROVIDE A HIGH ENERGY FLAME WITHIN SAID COMBUSTION CHAMBER FOR INITIATING COMBUSTION OF SAID PRIMARY FUEL, (F) SAID FLAME PRODUCER COMPRISING A PYROTECHNIC FLARE SUPPORTED WITHIN SAID COMBUSTION CHAMBER AND HAVING A CHARGE OF HIGH ENERGY SOLID ROCKET FUEL INCLUDING AN OXIDIZER AND MEANS IGNITING SAID ROCKET FUEL, (G) SAID FLARE COMPRISING A METAL SLEEVE SECURED TO SAID HOUSING STRUCTURE AND HAVING AN OPEN END DISPOSED ADJACENT SAID COMBUSTOR STRUCTURE AND EXTERIORLY OF SAID COMBUSTION CHAMBER, (H) SAID FLARE FURTHER COMPRISING A COMBUSTIBLE TUBE CARRYING SAID CHARGE OF ROCKET FUEL AND CARRIED WITHIN SAID SLEEVE AND EXTENDING FROM SAID OPEN END THROUGH SAID COMBUSTOR STRUCTURE AND INTO SAID COMBUSTION CHAMBER, WHEREBY SAID SLEEVE IS DISPOSED SUBSTANTIALLY ENTIRELY EXTERIORLY OF SAID COMBUSTION CHAMBER AND SAID COMBUSTIBLE TUBE IS DISPOSED SUBSTANTIALLY ENTIRELY INSIDE SAID COMBUSTION CHAMBER.
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Cited By (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3240016A (en) * 1960-03-16 1966-03-15 Nathan C Price Turbo-jet powerplant
US3690096A (en) * 1969-10-10 1972-09-12 Harry Munby Igniter arrangement for a gas turbine engine
US4040251A (en) * 1975-06-04 1977-08-09 Northern Research And Engineering Corporation Gas turbine combustion chamber arrangement
US4545196A (en) * 1982-07-22 1985-10-08 The Garrett Corporation Variable geometry combustor apparatus
US20040025495A1 (en) * 2000-09-05 2004-02-12 Dev Sudarshan Paul Nested core gas turbine engine
US20050076650A1 (en) * 2003-10-08 2005-04-14 Rodolphe Dudebout Auxiliary power unit having a rotary fuel slinger
US7937946B1 (en) 2005-12-21 2011-05-10 Florida Turbine Technologies, Inc. Small gas turbine engine with lubricated bearings
US20110219779A1 (en) * 2010-03-11 2011-09-15 Honeywell International Inc. Low emission combustion systems and methods for gas turbine engines
US8479492B2 (en) 2011-03-25 2013-07-09 Pratt & Whitney Canada Corp. Hybrid slinger combustion system

Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
GB627722A (en) * 1946-11-01 1949-08-15 John Richard Marsden Improvements in or relating to gas-turbine-engines
US2518882A (en) * 1948-12-04 1950-08-15 Daniel And Florence Guggenheim Combined igniting and fuel feed control mechanism for a thrust unit
US2561670A (en) * 1945-07-30 1951-07-24 Aerojet Engineering Corp Ignitor
US2595759A (en) * 1948-11-30 1952-05-06 Gen Electric Atomizing nozzle for spraying viscous liquids
GB714800A (en) * 1952-05-22 1954-09-01 France Etat Improvements in or relating to explosively-operated ignition devices for jet engines
US2760340A (en) * 1954-08-30 1956-08-28 Clifford E Seglem Igniter and combustion apparatus
US2784553A (en) * 1954-06-10 1957-03-12 Corso Serafino M De Combustion conduit and igniter structure
US2856755A (en) * 1953-10-19 1958-10-21 Szydlowski Joseph Combustion chamber with diverse combustion and diluent air paths
US2937501A (en) * 1955-07-13 1960-05-24 Sud Aviation Combustion devices for ram-jet engines

Patent Citations (9)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2561670A (en) * 1945-07-30 1951-07-24 Aerojet Engineering Corp Ignitor
GB627722A (en) * 1946-11-01 1949-08-15 John Richard Marsden Improvements in or relating to gas-turbine-engines
US2595759A (en) * 1948-11-30 1952-05-06 Gen Electric Atomizing nozzle for spraying viscous liquids
US2518882A (en) * 1948-12-04 1950-08-15 Daniel And Florence Guggenheim Combined igniting and fuel feed control mechanism for a thrust unit
GB714800A (en) * 1952-05-22 1954-09-01 France Etat Improvements in or relating to explosively-operated ignition devices for jet engines
US2856755A (en) * 1953-10-19 1958-10-21 Szydlowski Joseph Combustion chamber with diverse combustion and diluent air paths
US2784553A (en) * 1954-06-10 1957-03-12 Corso Serafino M De Combustion conduit and igniter structure
US2760340A (en) * 1954-08-30 1956-08-28 Clifford E Seglem Igniter and combustion apparatus
US2937501A (en) * 1955-07-13 1960-05-24 Sud Aviation Combustion devices for ram-jet engines

Cited By (16)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3240016A (en) * 1960-03-16 1966-03-15 Nathan C Price Turbo-jet powerplant
US3690096A (en) * 1969-10-10 1972-09-12 Harry Munby Igniter arrangement for a gas turbine engine
US4040251A (en) * 1975-06-04 1977-08-09 Northern Research And Engineering Corporation Gas turbine combustion chamber arrangement
US4545196A (en) * 1982-07-22 1985-10-08 The Garrett Corporation Variable geometry combustor apparatus
US20040025495A1 (en) * 2000-09-05 2004-02-12 Dev Sudarshan Paul Nested core gas turbine engine
US20070201974A1 (en) * 2000-09-05 2007-08-30 Dev Sudarshan P Nested core gas turbine engine
US6988357B2 (en) * 2000-09-05 2006-01-24 Sudarshan Paul Dev Nested core gas turbine engine
US7219490B2 (en) 2000-09-05 2007-05-22 D-Star Engineering Nested core gas turbine engine
US20070012026A1 (en) * 2000-09-05 2007-01-18 Dev Sudarshan P Nested core gas turbine engine
US7036321B2 (en) * 2003-10-08 2006-05-02 Honeywell International, Inc. Auxiliary power unit having a rotary fuel slinger
US20050076650A1 (en) * 2003-10-08 2005-04-14 Rodolphe Dudebout Auxiliary power unit having a rotary fuel slinger
US7937946B1 (en) 2005-12-21 2011-05-10 Florida Turbine Technologies, Inc. Small gas turbine engine with lubricated bearings
US20110219779A1 (en) * 2010-03-11 2011-09-15 Honeywell International Inc. Low emission combustion systems and methods for gas turbine engines
US8479492B2 (en) 2011-03-25 2013-07-09 Pratt & Whitney Canada Corp. Hybrid slinger combustion system
US8677731B2 (en) 2011-03-25 2014-03-25 Pratt & Whitney Canada Corp. Hybrid slinger combustion system
US9874148B2 (en) 2011-03-25 2018-01-23 Pratt & Whitney Canada Corp. Hybrid slinger combustion system

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