US3046731A - Flame stabilization in jet engines - Google Patents

Flame stabilization in jet engines Download PDF

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US3046731A
US3046731A US539116A US53911655A US3046731A US 3046731 A US3046731 A US 3046731A US 539116 A US539116 A US 539116A US 53911655 A US53911655 A US 53911655A US 3046731 A US3046731 A US 3046731A
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jet
fuel
air
flame
afterburner
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US539116A
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Cambel Ali Bulent
Allan B Schaffer
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EDWARD POHLMANN
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EDWARD POHLMANN
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    • FMECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23RGENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/16Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration with devices inside the flame tube or the combustion chamber to influence the air or gas flow
    • F23R3/18Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants
    • F23R3/24Flame stabilising means, e.g. flame holders for after-burners of jet-propulsion plants of the fluid-screen type

Description

y 1962 A. B. CAMBEL ETAL FLAME STABILIZATION IN JET ENGINES Filed Oct 7, 1955 0- Ohm u 02 km: Eat-.0:
0 g 8 (ScL-l) AJJOO'IBA :L-IOMO'IB INVENTORS: ALl BULENT CAMBEL ALLAN B. CHAFFER 3,46,73l Patented July 31, 1962 Fire 3,046,731 FLAME STABILIZATION IN JET ENGINES Ali Bulent Cambel, 181i Hinman Ave., Evanston, ill,
and Allan B. Schaifer, Box 495, Rte. 3, Salem, Va, as-
signors of twenty-five percent to Edward Pohlrnalm,
East Orange, NJ.
Filed Oct. 7, 1955, Ser. No. 539,116 11 Claims. (Cl. oil-35.6)
invention relates to flame stabilization in jet engines and more particularly to the stabilization of the flame in the afterburner of a jet engine such as is used in airplanes. The term jet engine as usedherein is intended to include both turbojets and ramjets but the invention will be described particularly with respect to turbojets.
In a turbojet engine intake air is compressed up to say twelve times its normal density and from the compressor section flows into a burning section; Here fuel is constantly pumped into the air stream under great pressure, atomized through nozzles and constantly burned. The burning causes great expansion which speeds up the fiow of the gases. The gases pass through one or more turbine wheels and then out through the nozzle. The turbine wheels in turn operate the compressor. A little more than one-half the energy in the expelled gases is taken out of them in the form of rotating motion by the turbine wheel or wheels and passed back to the compressor to do the initial work of increasing the intake density. All of the rest of the energy which is expelled through the nozzle provides the thrust for the m'rplane in which the jet engine is mounted. In order to provide an increased amount of thrust, it is the usual practice in high powered turbojet engines, such as are used in airplanes, to burn a part of the fuel in a primary burner and another part in an afterburner. The primary burner is normally located as previously mentioned between the compressor and the turbine. between the turbine and the nozzle, The nozzle is normally of the variable-area type when an afterburner is employed.
In the primary burner the effective velocity of the approach gases is relatively slow, say 40 to '50 feet per second. In the afterburner, on the other hand, the velocity of the approach gases is usually within the range of 300 to 70 feet per second. One of the difficulties encountered, however, is the tendency for the flame in the afterburner to blow out due to the high approach velocity. Elforts have heretofore been made to stabilize the flame in the afterburner by the use of mechanical devices which are usually called V-gutters.
One of the objects of the present invention is to provide a method of stabilizing a high velocity flame by a relatively simple device which does not involve the use of mechanical flanieholders such as V-gutters or the like.
Another object of the invention is to provide a jet engine construction and method of operation which makes it possible to operate an 'afterburner at a high approach velocity and with a wide range for the stability limits.
Another object of the invention is to provide a jet engine capable of stabilizing approach velocities of 300 to 700 feet per second at temperatures in excess of 1000 F. and at afterburner pressures varying from two atmospheres at sea level to a fraction of an atmosphere at high altitudes.
Another object of the invention is to provide a flame stabilizer with a low dry loss, that is, with a low pressure drop across the flameholder when the afterburner is not in use.
Another object of the invention is to provide a flame- The afterburner is located holder that will allow shorter afterburners by producing high flame spreading together with low dry loss.
Other objects and advantages of the invention will be apparent by reference to the following description in conjunction with the accompanying drawing in which:
FIG. 1 illustrates diagrammatically a turbojet engine;
FIG. 2 is a sectional view taken along the line 22 of FIG. 1;
FIG. 3 is a modified form of FIG. 2;
7 FIG. 4 is an enlarged detail of one form of opposing jet manifold used in the structure of FIG. 2 or FIG. 3;
FIG. 5 is an enlarged detail of a modified type of opposing jet manifold; and
FIG. 6 is a graph showing performance curves with varying types of opposing jets used as flame stabilizers.
In accordance with the invention it has been found that a high velocity flame in a jet engine can be stabilized by means of a high velocity opposing jet of combustionsustaining media such as air, oxygen, or vaporized fuelair mixtures. The stabilization is accomplished by the creation of a low velocity stagnation region into which are fed primary media, jet media, and hot burned gases entrained by the jet stream. The mass flow of jet media is of the order of magnitude of 4% to 2%, preferably around 1%, of the primary mass flow. For the customary aircraft application an opposing air jet can be employed and a minute quantity of secondary fuel can be introduced into the jet stream so that the system exhibits its peak performance at the normal operating fuel-air ratio of the primary mixture.
Referring to the drawings, in FIG. 1 la turbojet englue is diagrammatically illustrated and comprises a compressor section 1, a primary stage combustion chamber 2., a plurality of fuel inlets 3 adapted to spray or atomize the fuel into the combustion chamber 2, and a turbine 4 which drives the compressor 1 through the drive shaft 5. The afterburner generally indicated at 6 has guide cone '7 and a tail pipe 8 which terminates in a nozzle or opening 9. A plurality of fuel nozzles 10 are disposed in the afterburner 6 for the combustion of fuel to increase the temperature of the hot gases passing through the tail pipe 8.
In accordance with the present invention a plurality of opposing orifices 11 are provided to stabilize the flame in the afterburner 6. The orifices 11 are preferably mounted concentrically as shown in FIG. 2 or in a grid pattern as shown in FIG. 3 but can be mounted in any suitable manner within the afterburner 6. In the concentric arrangement shown in FIG. 2, the combustion-sustaining media such as air, is delivered through a passageway 12 from a point 13 just back of the compressor l to a point 14 where it enters a spider-like arrangement of tubes 15, generally shown in FIG. 2. The passageway 12 may consist of a passageway which extends entirely around the jet engine fuselage or it can consist of a plurality of tubes connected with the tubes of the spider-like opposing jet structure 15 at the points 14. Similarly, in FIG. 3 the passageway .12 connects with the grid-type opposing jet structure 15 at any of the points 14'. Moreover, as in the case of the structure shown in FIG. 2, the passageway 12' instead of surrounding the entire fuselage can consist of individual pipes connecting the opposing jet-containing elements with a source of combustion-sustaining gas.
Tubes 16 and 16, respectively, shown in FIGS. 2 and 3, which carry the orifices 11 for the opposing jets preferably have the cross-sectional shape of an airfoil as shown in FIG. *4. (if the opposing jet stream is composed of a mixture of air and fuel it is sometimes desirable to introduce the fuel through a separate line 17 to be mixed with the air at the orifice 1 1 as shown in FIG. 5.
In FIG. 5 the opposing jet is created in part by the bent tube 17.
The diameter and the number of the orifices or jets 11 will vary depending upon the particular application, that is, upon the size of the afterburner 6 and the velocity, pressure and temperature of the gases therein. The opposing jets can be subsonic, sonic, or supersonic, but a sonic jet is most convenient. The optimum angle of attack for the jets to the primary stream is 180, although variations of a few degrees generally not more than 10, from 180 are permissible. The use of an airfoil shape in the jet manifolds 16 as shown in FIG. 4 reduces the dry loss. Secondary fuel can be introduced by several methods: (1) by premixing with the jet air; (2) by spraying into the nose of the flame from an upstream position; or (3) by spraying from a small tube concentric within each air jet hole as shown in FIG. 5. Suitable automatic control apparatus is employed to meter the secondary fuel in the desired proportion to the jet air as the latter varies. Guide vanes can be employed upstream of the stabilized flames in order to reduce the effects of swirl of the primary mixture; that is, the effects of the varying angle of attack of primary mixture. High heating of the jet media can be effected, if desired, by ducting the media through the flame.
As an example of an application, an opposing air jet with a supply pressure of pounds per square inch absolute issuing from a A inch hole into 1% inch diameter combustion chamber supplied with a non-preheated propane-air primary mixture can stabilize approach velocities as high as 340 feet per second with peak performance occurring at a fuel-air ratio of 1.4 times the stoichiometric value. When secondary fuel is introduced in the proper amount the peak performance can be made to occur at virtually any Value less than 1.4, or, of particular significance, at the stoichiometric value of the fuel-air ratio.
These results are shown in FIG. 6 which is a graph showing the performance of a room temperature (70 F.) air jet, a hot air jet (supplied at about 1000 F.) and a room temperature (70 F.) stoichiometric jet of propane and air. In the tests which were made to obtain the data employed in preparing the graph of FIG. 6, the primary fuel employed was propane. The opposing jets were supplied at a pressure of 70 pounds per square inch gauge. The diameter of the jet orifice was inch. The primary air was supplied at a temperature of 70 F. The flame was burned inside a chamber having an inside diameter of 1.77 inches. The pressure at the exhaust end of the chamber was one atmosphere. The length of the flame in the chamber was 5 inches and the tube carrying the jet orifice extended 3 inches into the exhaust end of the chamber. The term approach velocity indicates the velocity in the gases approaching the opposing jet in feet per second. The graphs on FIG. 6 represent the blowoif velocity, that is to say, the approach velocity of the gases when the flame blows out because under the specific conditions the approach velocity is too high. The equivalence ratio is the ratio of the actual fuel-air ratio to the stoichiometric fuel-air ratio. The stoichiomet-ric fuel-air ratio is the ratio of fuel to air where the amount of air is just sufiicient to burn all of the fuel in the mixture. Thus, when the equivalence ratio is greater than 1, the mixture is rich and when it is less than 1, the mixture is lean. It will be observed that the air jet results in a higher blowoff velocity with a rich mixture. When the equivalence ratioof the mixture is 1, the optimum blowoff velocity is obtained by employing an opposing jet in which the equivalence ratio of propane and air is stoichiometric. Another point to be observed in connection with the use of the air jet as a flameholder is the relatively wider range of equivalence ratios where the mixture is rich.
One of the advantages of the opposing jet flameholder is that it is possible to provide a relatively large number of jet openings to serve as flameholders while blocking only a small fraction of the area of the afterburner as compared to that which is blocked by a standard type of mechanical flameholder such as a V-gutter. Furthermore, it is possible to vary the effect of the opposing jet by increasing or decreasing the pressure of the gases used in the opposing jet. The exact pressure used in the opposing jet flameholder will vary, of course, depending upon the approach velocity of the main gas stream and the number of jet flameholders employed. As a consequence of the small area of the afterburner which is blocked by the opposing jet structure, there is much less pressure drop across the jet flameholder when it is not in use and the thrust can be increased by 3% to 5%. if the opposing jet flameholders are supplied with air, the air can come from the compressor and the small amount of air bled from the compressor does not represent a substantial loss.
As illustrated with respect to FIG. 6, the performance curves for an opposing jet flameholder can be shifted (with respect to the equivalence ratio axis) by varying the equivalence ratio of the jet media. The shift will be toward lean ratios for increasing amounts of fuel in the opposing jet. This will allow leaner mixtures to be burned in the afterburner thereby conserving fuel.
For high altitude flight (say, greater than 100,000 feet) where stabilization is ordinarily diflicult, an opposing jet enriched with oxygen will considerably facilitate stabilization.
The invention also makes it possible to eliminate certain combustion instabilities associated with mechanical flameholders which are believed to be due to vortex shedding from the physical boundary of the V-gutter.
In general, therefore, the invention is useful in stabilizing a flame at approach velocities, temperatures and pressures encountered in afterburners of turbojets and in ramjets. The stabilization is effected with significantly less dry loss than with V-gutters of the size and blockage now employed in aircraft and with less pressure drop across a stabilized flame as compared with a V-gutter, thereby allowing greater thrust when the afterburner is not in use and also greater thrust when the afterburner is in use.
The invention provides high flame spreading, together with low dry loss so that a plurality of opposing jet air fiameholders can be employed and can be closely spaced.
The result is that the flame will cover the entire crosssection of the afterburner in a short distance, thereby permitting the afterburner to 'be shorter. At the present time, afterburners are about 13 feet long and hence it will be appreciated that a shorter afterburner is highly desirable.
The invention also provides a flexible flameholding system which, by adjusting the flow of secondary fuel, can be made to stabilize a flame at virtually any level of thrust augmentation desired. Hence, afterburner modulation can be readily achieved, that is to say, the thrust level of the afterburner can be raised gradually.
The invention also eliminates instabilities such as screec which are associated with vortex shedding from V-gutters.
The invention also makes it possible to provide a mechanism for the ready utilization of high energy fuels. These fuels can be used as primary or secondary fuels by premixing or concentric injection with the opposed air jet.
-he invention is hereby claimed as follows:
1. A jet engine comprising an afterburner provided with primary fuel mixture discharge means, an exit opening for exhaust gases and means in said afterburner downstream of said primary fuel mixture discharge means and upstream of said exit opening providing a jet of a combustion sustaining fluid in a direction opposed to the flow of the fuel mixture from said fuel discharge means and in an amount corresponding to a minor proportion by volume compared to the volume of the primary flow of fuel mixture, said fuel mixture providing a flame originating upstream of said jet means and said jet being adapted to stabilize the flame in said afterburner.
2. A jet engine as claimed in claim 1 in which said com bustion sustaining fluid is air.
3. A jet engine as claimed in claim 1 in which said combustion sustaining fluid is oxygen.
4. A jet engine as claimed in claim 1 in which said combustion sustaining fluid is a mixture of air and fuel.
5. A jet engine comprising an afterburner, a primary fuel mixture discharge means in said afterburner, an exhaust exit from said afterburner through which products of combustion are exhausted, a plurality of orifices positioned in said afterburner between said primary fuel mixture discharge means and said exhaust exit, said orifices being adapted to discharge jets of a combustion sustaining fluid in a direction opposed to the primary flow of said fuel mixture and in an amount corresponding to a minor proportion by volume compared to the volume of the primary flow of fuel mixture, said fuel mixture providing a flame originating upstream of said orifices and said jet-s serving as the sole stabilization means for the flame produced by the combustion of said primary fuel mixture.
6. A method of stabilizing a flame in an afterburner of a jet engine which comprises forcing a combustion sustaining fluid through an orifice in a direction opposed to the primary flow of fuel mixture producing said flame, said fuel mixture providing a flame originating upstream of said orifice and said combustion sustaining fluid being a minor proportion by volume compared to the volume of said primary fuel mixture and being suflicient to stabilize said flame.
7, A method as claimed in claim 6 in which the mass flow of the fluid through said orifice is around tc 2% of the primary mass flow of the fuel mixture producing said flame.
8. A method as claimed in claim 6 in which said combastion sustaining fluid is air.
9. A method as claimed in claim 6 in which said combustion sustaining fluid is oxygen.
10. A method as claimed in claim 6 in which said combustion sustaining fluid is composed of a mixture of fuel and a combustion sustaining gas.
11. A method as claimed in claim 6 in which the primary fuel mixture is composed of air and fuel and the combustion sustaining fluid passed through said orifice is composed of a mixture of air and fuel having approximately the same equivalence ratio as the air and fuel in said primary fuel mixture.
References Cited in the file of this patent UNITED STATES PATENTS 2,440,491 Schwander Apr. 27, 1948 2,483,737 Parrish Oct. 4, 1949 2,517,015 Mock et al. Aug. 1, 1950 2,622,396 Clarke et al. Dec. 23, 1952 2,628,473 Frye Feb. 17, 1953 2,720,078 Day et a l. Oct. 11, 1955 2,734,341 Lovesey Feb. 14, 1956 2,771,743 Lovesey Nov. 27, 1956 FOREIGN PATENTS 1,085,458 France July 28, 1954 619,251 Great Britain Mar. 7, 1949 713,265 Great Britain Aug. 11, 1954 761.167 Great Britain Nov. 14 1956
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Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3295318A (en) * 1964-06-09 1967-01-03 Snecma Burner devices and combustion chambers therefor
US3300976A (en) * 1964-02-21 1967-01-31 Rolls Royce Combined guide vane and combustion equipment for bypass gas turbine engines
US3413810A (en) * 1965-05-15 1968-12-03 Bolkow Gmbh Fuel injection device for liquid fuel rocket engines
US3465525A (en) * 1966-03-25 1969-09-09 Rolls Royce Gas turbine bypass engines
US3600892A (en) * 1968-06-10 1971-08-24 Technology Uk Combustion devices
US3973390A (en) * 1974-12-18 1976-08-10 United Technologies Corporation Combustor employing serially staged pilot combustion, fuel vaporization, and primary combustion zones
US4455840A (en) * 1981-03-04 1984-06-26 Bbc Brown, Boveri & Company, Limited Ring combustion chamber with ring burner for gas turbines
US5867980A (en) * 1996-12-17 1999-02-09 General Electric Company Turbofan engine with a low pressure turbine driven supercharger in a bypass duct operated by a fuel rich combustor and an afterburner
EP2584267A3 (en) * 2011-10-17 2017-11-15 General Electric Company Injector having mulitple fuel pegs
FR3094777A1 (en) * 2019-04-04 2020-10-09 Safran Aircraft Engines Main combustion chamber of a turbojet engine fitted with a grid downstream of its burners

Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2440491A (en) * 1935-03-21 1948-04-27 Des Vehicules Sev Soc Et Oil burner
GB619251A (en) * 1946-11-27 1949-03-07 Donald Louis Mordell Improvements relating to combustion-equipment
US2483737A (en) * 1943-07-10 1949-10-04 Stewart Warner Corp Internal-combustion burner for heaters
US2517015A (en) * 1945-05-16 1950-08-01 Bendix Aviat Corp Combustion chamber with shielded fuel nozzle
US2622396A (en) * 1948-11-26 1952-12-23 Lucas Ltd Joseph Fuel vaporizing apparatus for an afterburner
US2628473A (en) * 1948-05-03 1953-02-17 Frye Jack Stationary power plant having radially and axially displaced jet engines
GB713265A (en) * 1951-08-10 1954-08-11 Rolls Royce Improvements relating to combustion equipment for gas turbine engines
FR1085458A (en) * 1953-06-27 1955-02-02 Snecma Improvements to combustion devices
US2720078A (en) * 1948-03-01 1955-10-11 Solar Aircraft Co Burner for use in high velocity ducts
US2734341A (en) * 1956-02-14 Reheating turbine exhaust gases
GB761167A (en) * 1953-06-27 1956-11-14 Snecma Flame spreading device for combustion systems
US2771743A (en) * 1951-08-10 1956-11-27 Rolls Royce Gas-turbine engine with reheat combustion equipment

Patent Citations (12)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US2734341A (en) * 1956-02-14 Reheating turbine exhaust gases
US2440491A (en) * 1935-03-21 1948-04-27 Des Vehicules Sev Soc Et Oil burner
US2483737A (en) * 1943-07-10 1949-10-04 Stewart Warner Corp Internal-combustion burner for heaters
US2517015A (en) * 1945-05-16 1950-08-01 Bendix Aviat Corp Combustion chamber with shielded fuel nozzle
GB619251A (en) * 1946-11-27 1949-03-07 Donald Louis Mordell Improvements relating to combustion-equipment
US2720078A (en) * 1948-03-01 1955-10-11 Solar Aircraft Co Burner for use in high velocity ducts
US2628473A (en) * 1948-05-03 1953-02-17 Frye Jack Stationary power plant having radially and axially displaced jet engines
US2622396A (en) * 1948-11-26 1952-12-23 Lucas Ltd Joseph Fuel vaporizing apparatus for an afterburner
GB713265A (en) * 1951-08-10 1954-08-11 Rolls Royce Improvements relating to combustion equipment for gas turbine engines
US2771743A (en) * 1951-08-10 1956-11-27 Rolls Royce Gas-turbine engine with reheat combustion equipment
FR1085458A (en) * 1953-06-27 1955-02-02 Snecma Improvements to combustion devices
GB761167A (en) * 1953-06-27 1956-11-14 Snecma Flame spreading device for combustion systems

Cited By (10)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US3300976A (en) * 1964-02-21 1967-01-31 Rolls Royce Combined guide vane and combustion equipment for bypass gas turbine engines
US3295318A (en) * 1964-06-09 1967-01-03 Snecma Burner devices and combustion chambers therefor
US3413810A (en) * 1965-05-15 1968-12-03 Bolkow Gmbh Fuel injection device for liquid fuel rocket engines
US3465525A (en) * 1966-03-25 1969-09-09 Rolls Royce Gas turbine bypass engines
US3600892A (en) * 1968-06-10 1971-08-24 Technology Uk Combustion devices
US3973390A (en) * 1974-12-18 1976-08-10 United Technologies Corporation Combustor employing serially staged pilot combustion, fuel vaporization, and primary combustion zones
US4455840A (en) * 1981-03-04 1984-06-26 Bbc Brown, Boveri & Company, Limited Ring combustion chamber with ring burner for gas turbines
US5867980A (en) * 1996-12-17 1999-02-09 General Electric Company Turbofan engine with a low pressure turbine driven supercharger in a bypass duct operated by a fuel rich combustor and an afterburner
EP2584267A3 (en) * 2011-10-17 2017-11-15 General Electric Company Injector having mulitple fuel pegs
FR3094777A1 (en) * 2019-04-04 2020-10-09 Safran Aircraft Engines Main combustion chamber of a turbojet engine fitted with a grid downstream of its burners

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