US3859787A - Combustion apparatus - Google Patents

Combustion apparatus Download PDF

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Publication number
US3859787A
US3859787A US439648A US43964874A US3859787A US 3859787 A US3859787 A US 3859787A US 439648 A US439648 A US 439648A US 43964874 A US43964874 A US 43964874A US 3859787 A US3859787 A US 3859787A
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United States
Prior art keywords
prechamber
air
combustion
fuel
liner
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US439648A
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English (en)
Inventor
Robert D Anderson
Dennis L Troth
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Motors Liquidation Co
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General Motors Corp
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Priority to US439648A priority Critical patent/US3859787A/en
Application filed by General Motors Corp filed Critical General Motors Corp
Priority to CA212,069A priority patent/CA1013153A/en
Priority to GB5374074A priority patent/GB1467499A/en
Priority to DE2460709A priority patent/DE2460709C2/de
Publication of US3859787A publication Critical patent/US3859787A/en
Application granted granted Critical
Priority to IT47881/75A priority patent/IT1026493B/it
Priority to SE7500960A priority patent/SE393839B/xx
Priority to JP50014046A priority patent/JPS5825931B2/ja
Priority to FR7503466A priority patent/FR2259989B1/fr
Anticipated expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/02—Continuous combustion chambers using liquid or gaseous fuel characterised by the air-flow or gas-flow configuration
    • F23R3/26—Controlling the air flow
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, 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
    • F02C7/12—Cooling of plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, 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
    • F02C7/12—Cooling of plants
    • F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/28—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply
    • F23R3/30—Continuous combustion chambers using liquid or gaseous fuel characterised by the fuel supply comprising fuel prevapourising devices
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2250/00—Geometry
    • F05D2250/40—Movement of components
    • F05D2250/41—Movement of components with one degree of freedom
    • F05D2250/411—Movement of components with one degree of freedom in rotation

Definitions

  • ABSTRACT A combustion apparatus for gas turbine engines particularly adapted to reduce emissions to meet automotive requirements.
  • the fuel is laid on the wall of a cylindrical prechamber and evaporated from the wall by combustion air which is introduced through a swirler at the upstream end of the prechamber.
  • the inner surface of the prechamber is artificially roughened by a grid of grooves to improve fuel evaporation.
  • the fuel is laid on the wall from an annular manifold extending around the upstream end of the prechamber through tangential orifices leading from the manifold into the interior of the prechamber.
  • the fuel manifold is cooled and shielded from heat by an air jacket. More air enters through entrance ports distributed around the prechamber toward its downstream end.
  • the re sulting lean fuel-air mixture is delivered past an annular flow dam at the outlet of the: prechamber into a combustion or reaction zone which is abruptly renlarged from the prechamber.
  • the structure causes turbulent flow, recirculation, and good mixing in the reaction zone.
  • a dilution zone downstream at the reaction zone has a circumferential array of dilution air ports which are of such shape as to be varied nonlinearly in area by a sliding ring valve.
  • the sliding ring valve is coupled to a second sliding ring valve which varies the area of the air entrance ports in the pre chamber in reverse sense to the dilution air ports.
  • a pilot fuel nozzle to aid in cold starts is mounted at the upstream end of the prechamber.
  • COMBUSTION APPARATUS This invention is directed to combustion chambers of a type suitable for use with gas turbine engines. It is particularly directed to combustion chamber structures adapted to insure complete combustion over relatively wide ranges of air and fuel flow and to minimize discharge of incompletely burned fuel and generation of oxides of nitrogen.
  • the combustion chamber in which the present inven tion is embodied has been tested and found to be capable of meeting the 1977 emission standards in a suitable vehicle installation.
  • the invention which is the subject of this patent application is one of a number of improvements to a gas turbine combustion chamber of a type generally known in the prior art. These refinements cooperate to bring emissions down to the prescribed level.
  • the invention is described here in terms of its preferred embodiment, which includes others of these improvements.
  • the particular subject matter to which this application is directed is an improved fuel prevaporization portion of the combustion apparatus, with particular regard to improving the evaporation of the fuel from a surface by hot air flowing over the surface. This is achieved by texturing or grooving the surface to provide shallow channels for the fuel, with a grid of bosses between the channels.
  • the principal objects of this invention are to provide an improved combustion apparatus suitable for automotive use, to provide a combustion chamber having exceptionally clean exhaust, and to provide an improved arrangement for evaporating fuel from a surface and mixing it with air flowing over the surface.
  • FIG. 1 is a longitudinal view of a combustion apparatus for a gas turbine engine, with parts cut away and in section.
  • FIG. 2 is a longitudinal sectional view of the combustion liner.
  • FIG. 3 is an upstream end view of the combustion liner taken on the plane indicated by the line 33 in FIG. 2.
  • FIG. 4 is a cross sectional view of the prechamber taken on the plane indicated by the line 4-4 in FIG. 2.
  • FIG. 5 is a fragmentary longitudinal sectional view of the prechamber.
  • FIG. 6 is a detailed sectional view taken on the plane indicated by the line 66 in FIG. 5.
  • FIG. 7 is a fragmentary view taken on the plane indicated by the line 77 in FIG. 5.
  • FIG. 8 is a fragmentary view of the interior of the prechamber illustrating a textured surface.
  • FIG. 9 is a cross-section of the same taken in the plane indicated by the line 99 in FIG. 8.
  • FIG. 10 is a fragmentary exterior view of the prechamber wall.
  • FIG. 11 is a plot of airflow distribution.
  • FIG. 12 is a partial elevation view illustrating an alternative prechamber arrangement.
  • a gas turbine engine 2 includes an engine case 3. Further details of the engine are not shown or described, since they are immaterial to an understanding of the present invention. By way of background, however, the engine may be a regenerative gas turbine of the general nature of those described in US. patents to Collman et al. Pat. No. 3,077,074, Feb. l2, 1963;Collman et al. Pat. No. 3,267,674, Aug. 23, I966; and Bell Pat. No. 3,490,746, Jan. 20, I970.
  • the engine case 3 forms part of an outer casing 4 of the combustion apparatus which also includes a cylindrical housing 6 bolted to the engine case.
  • the engine compressor (not illustrated) delivers compressed air which is heated in a regenerator (not illustrated) on its way into the combustion apparatus casing 4.
  • Housing 6 terminates in a flange 7 to which is fixed a continuous outer ring 8 of a combustion liner support spider 10 which provides part of the support for a combustion liner II in which the invention is embodied.
  • Ring 8 is fixed to flange 7 by circumferentially spaced countersunk screws 12.
  • a combustion chamber cover 14 which overlies the ring 8 is fixed to the flange 7 by a ring of bolts 15 which extend through the ring and flange.
  • the housing 6 and cover 14 are lined with thermal insulating material 16.
  • the combustion liner 11 in its preferred form is of circular cross section and is bounded by walls 18.
  • the liner wall includes a first prechamber or fuel vaporizing portion 19 which extends to an abrupt radial enlargement defined by a substantially radially outwardly extending wall portion 20 which is integral with and continues into a cylindrical wall portion 22.
  • the wall portion 19 encloses a fuel vaporizing zone of the combustion apparatus and the wall 22 portion encloses a reaction zone 23 and a dilution zone 24.
  • Wall 22 terminates in an. outlet 25 for combustion products at the downstream end of the combustion liner. As shown in FIG. 1, the outlet end may be inserted into a combustion products duct 26 leading to the turbine (not shown). This supports the downstream end of the liner.
  • fuel is evaporated and the fuel and air are mixed in a prechamber 27 enclosed by wall section 19.
  • the fuel and air react, or combustion takes place, in the reaction zone 23 and additional air is introduced and mixed with combustion products in the dilution zone 24 to provide the ultimate mixture of combustion products to drive the turbine of the gas turbine engine.
  • a swirler 30 comprising an annular cascade of vanes 31 (see also FIGS. 3 and These vanes extend from an outer ring 32 to a central sleeve 34, the latter extending forwardly from the swirler 30.
  • the vanes of the swirler are set at an angle' of 75 to a plane extending axially of the combustion apparatus so as to impart a strong swirl component to air entering the liner at this point from the outer casing 4.
  • the outer ring 32 is welded or brazed to a prechamber forward wall portion 35. Wall section 35 is piloted within and fixed to the forward end of a rear prechamber wall portion 36.
  • Wall portion 36 is of relatively heavy stock, about A to 5/16 inch in thickness.
  • the downstream end of wall portion 36 is welded to the radially extending portion of the main combustion chamber wall, these parts being concentric.
  • Wall 20 extends radially inwardly from interior surface of wall portion 36 to provide a flow dam 38.
  • the remainder of the combustion liner wall is cylindrical and integral with the portion 20.
  • a sheet metal ring 39 extending over the forward portion of the prechamber has an inwardly extending flange 40 which is welded to the forward edge of wall portion 35.
  • This ring 39 provides for connecting the forward end of the combustion liner to the support spider 10.
  • the spider includes arms connecting outer ring 8 to an inner ring 42 (see also FIG. 1) which is suitably fixed or attached to the ring 39 of the liner.
  • the hot compressed air forced through swirler 30 will flow with a strong tangential component over the inner surface of wall portions 35 and 36 because of centrifugal force and will tend to scour these walls. In so doing, it vaporizes and picks up liquid hydrocarbon fuel which is fed to the inner surface of the prechamber just downstream of swirler 30.
  • the fuel is introduced from a manifold 46 (see FIGS. 2, 5, 6, and 7) which is a ring of semicircular section extending entirely around the outer surface of wall portion 35. Fuel is delivered to this manifold through a fuel inlet tube 47 which extends into the combustion apparatus from a suitable fitting for connection to an external source of supply (not illustrated).
  • Manifold 46 is enclosed within a cooling air jacket 48, likewise of semicircular cross section and likewise welded to the outer surface of wall section 35. Cooling air from a suitable source, for example from the compressor of the engine upstream of the regenerator, is supplied to the tube 48 through a cooling air pipe 50 which surrounds the fuel tube 47.
  • the cooling air jacket extends almost entirely around the prechamber,
  • the gap in the circumference ofthe tube is closed adjacent the inlet pipe by a semiannular blocking plate 51 brazed or welded to tubes 46 and 48 and to the wall portion 35.
  • Air introduced through tube 50 thus circulates over the fuel manifold to an outlet at 52 at the other end of the cooling air jacket. This circulation of air is to prevent boiling of the fuel under certain conditions of operation such as upon cutback of fuel with a hot engine, or during idling operation.
  • the support ring 39 also shields the fuel manifold and the cooling air jacket to some extent from heat which may be radiated from hot engine components near the flame tube.
  • Fuel supplied to the manifold 46 through tube 47 is laid on the interior of the prechamber wall through sixteen equally spaced main fuel ports or orifices 54. These ports are 0.0 l 3 inch in diameter and make about a 15 angle with the outer surface of the wall so that the fuel is squirted onto the inner surface of the wall rather than into the air flowing through the swirler.
  • the fuel is supplied at low pressure, the preferred maximum pressure drop through ports 54 being about 20 psi.
  • the current of air flowing through the swirler blows the introduced fuel along the inner surface of the prechamber wall portions 35 and 36, and the hot rapidly moving air heats and vaporizes and mixes with the fuel before entry into the reaction zone 23.
  • FIG. 8 is a view looking at the interior of the prechamber wall and FIG. 9 is a cross section of the same. The surface is relieved to provide a grid of two intersecting sets of small grooves 56 which leave between them small substantially rectangular bosses 55.
  • This sort of textured surface may most readily be achieved by coating the areas which provide the bosses 55 with a suitable resist and then etching the surface to the desired depth.
  • the resist may be applied by a photographic process, as is well understood.
  • the center to center spacing of adjacent grooves of each set is approximately 0.05 inch and the grooves are about 0.003 deep.
  • the width of each groove is about the same as the width of the bosses between the grooves.
  • Orientation of the grooves is preferably at about a 45 angle to the axial direction through the prechamber so that the fuel introduced into the inner wall may flow downstream of the prechamber under the influence of the air stream through the channels defined by the helically extending grooves 56.
  • burning of a lean mixture in the reaction zone 23 is preferable from the standpoint of clean exhaust to burning of a nearer to stoichiometric mixture. It is found desirable to introduce some air beyond that introduced by the swirler 30 to further mix with and dilute the fuel-air mixture prior to the initiation of combustion. This is effected by a set of air entrances distributed around the prechamber, preferably about three-fourths of its length from the upstream end to the downstream end.
  • the presently preferred structure for introduction of additional air introduces the air with radially inward and tangential components of movement and no significant axial component.
  • Equivalence ratio will be understood to mean the ratio of the actual weight ratio of fuel to air to the stoichiometric ratio of fuel to air. This is accomplished effectively by varying the ratio between the quantity of air flowing into the reaction zone from the prechamber to that introduced through dilution ports in the dilution zone 24 as the ratio of total airflow to fuel flow varies.
  • the wall portion 36 is made in two coaxial abutting sections fixed together, an upstream section 58 and a downstream section 59.
  • the air entrance means is defined by an annular array of slots 60 machined in the downstream edge of upstream section 58. Of course, they could be machined in the upstream portion of section 59 if the joint between the two sections is suitably located. It will be seen from FIGS. 4 and that slots 60 enter the chamber at a considerable angle to the radial, about 60 in the particular case, and are so oriented that the direction of swirl of air from these slots is the same as that imparted by the inlet swirler 30.
  • the outline of the slots is trapezoidal, the walls which bound the slots diverging from each other in the direction toward the upstream end of the prechamber.
  • the fragmentary view of FIG. 10 shows two such slots. In the total circumference there are preferably eighteen air entrance slots.
  • the wall section 58 also defines a radial port 62 through which an igniter 63 (FIG. 3), which may be similar to a spark plug, extends into the prechamber so as to light off the fuel.
  • an igniter 63 (FIG. 3)
  • the details of the igniter are immaterial, so it is not further described.
  • section 59 may bear three bosses 64. These provide a limit to movement of a flow modulating sleeve 66 slidably mounted on the exterior of the prechamber wall portion 36. As will be further described, this sleeve provides means for varying the flow of air through slots 60.
  • the dilution zone 24 is characterized by an array 70 of dilution air entrance ports, the effective area of which is varied by a ported axially slidable flow modulating sleeve 71.
  • the ports are of two sets alternating around the periphery of the liner.
  • One set is of ports 72, which are rectangular and of the least dimension axially of the liner.
  • the ports 74 of the second set which have an extension 75 toward the upstream end of the liner which is of smaller width circumferentially of the liner than the downstream portion of the ports 74.
  • the sleeve 71 is a simple cylinder with slightly flared ends and with a circumferential stiffening rib or ridge 78 near its upstream end. This sleeve is reciprocable on the outer surface of the liner, its travel being limited by two sets of small bosses 79 fixed to the outer surface of the liner, the bosses of each set being distributed 120 apart around the circumference of the liner.
  • the flow modulating sleeve 71 has two sets of ports each cooperating with one set of ports 72 or 74 of the liner wall. Rectangular ports 80 register with ports 72 of the wall and rectangular ports 82 register with ports 74. The length and width of the ports in the sleeve corresponds to the length and width of the corresponding port in the liner so that the liner ports can be fully opened at one setting of the sleeve.
  • One of the short ports 80 has two slots 83 extending axially of the liner at its upstream edge. These slots provide clearance for two pins 84 which extend outwardly from the liner and are welded to the liner wall.
  • pins serve to locate the sleeve 71 circumferentially of the liner and obviate any tendency for the sleeve to rotate around the axis of the liner as it is moved back and forth to vary the dilution air flow.
  • the sleeve 71 is moved by three axially movable rods 86 (FIG. 1) which are coupled by pins to webs 87 extending radially from the sleeve 71.
  • Rods 86 extend through guides 88 in the combustion chamber cover 14 and through seals or glands 90 to a common actuator plate 91 to which the rods are fixed by nuts 92.
  • the plate 91 may be coupled through a rod 94 to any suitable acutating mechanism capable of sliding the sleeve 71 axially of the liner.
  • the flow modulating sleeve 66 which varies the flow area through the ports 60 of the prechamber is rigidly coupled to sleeve 71 for concurrent movement by the input device 94.
  • This interconnection includes three rods 95 extending axially of and distributed around the circumference of the liner. Each rod is fixed at an anchorage 96 to the sleeve 71. They are adjustably connected to sleeve 66. This connection is effected through arms 102 fixed to the sleeve 66 lapart and extending radially outwardly. The arms are stiffened by gussets 103 welded to the arms and through a base plate 104 to the sleeve 66.
  • each rod 95 is threaded and extends through a hole at the extremity of an arm 102.
  • the connection is completed by two sets of double nuts 106 which may be adjusted to trim the relative poisitions of the two flow controlling sleeves. It will be seen that the two sets of sleeves move so that, as the ports 60 in the prechamber open, the ports 70 in the dilution area close.
  • the downstream edge of sleeve 66 is notched as indicated at 106 in FIGS. 1 and 2. There is a notch aligned with each air entrance slot 60, the notches being V- shaped and 'having an included angle of about 70. These notches are slightly wider than the downstream end of slots 60. They provide a tapering rather than an abrupt opening or closing of the slots 60 as the sleeve 66 moves so that its rear edge passes the rear edge of the slots.
  • FIG. 11 illustrates the variation of air flow with movement of the flow controlling sleeves. Specifically, curve A of FIG. 11 shows the proportion of air entering the reaction zone through the swirler and the slots 60 to the total air admitted, which is this amount of air plus that entering through the dilution ports at 70. It will be noted that the reaction zone air flow increases from about 17 percent with the sleeves at their maximum downstream position to approximately 55 percent with the sleeves moved forwardly to provide maximum reaction air flow relative to dilution air flow.
  • the dilution air flow decreases from about 83 percent to about 45 percent of total flow over this range of movement of the sleeves. This makes it possible to provide adequate air flow in the reaction zone under high power conditions without having the reaction zone undesirably rich as the power level of the engine is decreased. At small fuel flows, as under engine idling conditions, a relatively small part of the air is required to provide the desired equivalence ratio of about 0.3 in the reaction zone.
  • the bend at 107 in curve A represents closing of the short dilution ports 72.
  • a pilot fuel nozzle 110 is mounted in the prechamber.
  • This nozzle is preferably of an air-atomizing type supplied with compressed air and fuel through tubes (not illustrated) which enter through a supporting structure including a ring 111 fixed within the sleeve 34 by cap screws 112.
  • This arrangement provides a suitable support for the fuel nozzle which includes a tubular extension 114 threadably coupled with the ring 111.
  • the details of the fuel nozzle are not material to the present invention.
  • the pilot fuel nozzle is provided for starting combustion, particularly when the engine is cold and therefore evaporation of fuel from the prechamber wall is not effective.
  • the pilot nozzle is turned off after normal operation has begun. Other starting expedients such as use of gaseous fuel may be employed, but are not considered as feasible as the use of the pilot nozzle.
  • a converging fairing 115 extends from the downstream and of sleeve 34 to the downstream end of the fuel nozzle 110 to provide a smooth transition of flow from the swirler 30 into the prechamber.
  • the air enters the combustion apparatus at about l,lOOF. and after passing through the swirler 30 flows over the inside of the prechamber wall, heating, evaporating, and mixing with the fuel introduced from the manifold 46 through the orifices 54.
  • This mixture of air and vaporized fuel is further mixed with additional combustion air which enters through the swirl ports 60 with swirl in the same direction as the air flowing rear wardly through the prechamber.
  • These two flows then mix to provide a rather lean fuel-air mixture, preferably with about three times the amount of air required for combustion; that is, three times the stoichiometric amount of air.
  • the swirling fuel-air mixture spills over the dam 38 and because of the swirl flows tangentially and radially outward to the outer wall 22 of the liner and then, because of the creation of a low pressure area along the axis of the combustion zone and prechamber, it flows in a more or less toroidal vortex with some upstream or recirculating flow along the axis of the liner. This flow may penetrate into the downstream part of the prechamber under some conditions of operation and in this case may also tend to heat the prechamber.
  • the dilution air admitted through the openings at tends to quench the heat of the combustion mixture which flows along the wall 22 toward the outlet 25.
  • the radially entering streams of air as they meet toward the axis tend to project some of the dilution air forwardly into the low pressure zone on the combustion chamber axis where this mixes with the recirculating combustion products to assist in cooling the combustion products at an early time, reducing duration of high temperature in the gas.
  • the lean combustion lowers the combustion temperature and the prompt quenching of the gas lowers the residence time at high temperature. Both of these effects serve to minimize formation of nitrogen oxide.
  • the burning of the fuel in a vaporized condition reduces conditions of local richness which will be found in the vicinity of atomized droplets of fuel and which tend to increase generation of nitrogen oxide.
  • combustion liner shown which is for a 225 horsepower gas turbine engine, is 6 /a inches in diameter and 15 inches long, and is shown in true porportion in FIG. 2. It is possible, of course, to vary the ratio of primary to dilution air by throttling one set of ports only.
  • the broken line curve B in FIG. 11 illustrates the relation of reaction zone air flow to total flow in an apparatus of the sort illustrated in which the change of flow is due only to the movement of the sleeve 71, the downstream air entrance of the prechamber being of fixed area. For a given structure, the variation is less and the pressure drop is greater if modulation is effected at only one set of ports. Since pressure drops are inimical to engine efficiency, there is good reason to modulate both sets of air ports.
  • FIG. 12 illustrates a variation of the combustion liner which may be in most respects essentially as shown in figures previously discussed.
  • the liner of FIG. 12 differs from that of FIG. 2 in the mode of introduction of air into the prechamber, and the portion of the combustion liner downstream of that illustrated in FIG. 12 may be as illustrated in FIG. 2.
  • the prechamber wall 120 is a sheet metal structure of constant thickness bearing the swirler 30 at its forward end and with the central plug 34, 115 at the center of the swirler to support the starting fuel nozzle.
  • the interior of the prechamber wall is preferably textured as previously described.
  • the arrangement for introduction of additional primary air toward the downstream end of the prechamber in this case is a ring of circular holes 122 spaced uniformly around the prechamber.
  • the holes are 1/8 inch diameter and there are 36 holes.
  • a slightly smaller percentage of the air was admitted through the holes 1122 than through the ports 60 of FIG. 2.
  • the swirler 30 was slightly more open, having the blades set at a 70 angle to the axial direction rather than 75. No variation of the ports 122 was provided. While this apparatus does not perform as cleanly as that described above, it is a relatively clean combustion apparatus that might well serve quite satisfactorily in various applications.
  • the textured surface of this invention is instrumental in assuring this result in a compact prechamber, and is easily incorporated in the apparatus.
  • a combustion apparatus adapted for use in a gas turbine engine characterized by substantially complete combustion of liquid hydrocarbon fuel and by a low output of nitrogen oxides, the apparatus comprising a combustion liner having a discharge outlet for combustion products at the downstream end of the liner; the liner having an upstream end and liner wall means extending from the upstream end to the downstream end, the wall means enclosing, in sequence from the upstream end to the downstream end, a prechamber, a reaction zone, and a dilution zone; the prechamber including air entrance means at its upstream end effec' tive to direct combustion air with substantial velocity downstream over the inner surface of the prechamber wall means, and liquid fuel introduction means downstream of the said air entrance means disposed to lay a film of liquid fuel on the said inner surface for evaporation by and mixture in the prechamber with the said combustion air, wherein the improvement comprises a textured configuration of the said prechamber wall means inner surface defined by a grid of two sets of intersecting grooves in the surface and bosses rising between
  • a combustion apparatus adapted for use in a gas turbine engine characterized by substantially complete combustion of liquid hydrocarbon fuel and by a low output of nitrogen oxides, the apparatus comprising a combustion liner having a discharge outlet for combustion products at the downstream end of the liner; the liner having an upstream end and liner wall means extending from the upstream end to the downstream end,
  • the wall means enclosing, in sequence from the stream end to the downstream end, a prechamber, a reaction zone, and a dilution zone;
  • the prechamber including air entrance means defined by swirler means at its upstream end effective to direct combustion air with a substantial transverse velocity component downstream over the inner surface of the prechamber wall means, and liquid fuel introduction means downstream of the said air entrance means disposed to lay a film of liquid fuel on the said inner surface for evaporation by and mixture in the prechamber with the said combustion air
  • the improvement comprises a textured configuration of the said prechamber wall means inner surface defined by a grid of two sets of intersecting grooves in the surface and generally rectangular bosses rising between adjacent grooves, adapted to allow flow of the liquid fuel through the grooves and evaporation of the fuel from within the grooves by the flow of air over the said surface, the grooves being disposed at roughly a 45 angle to the axial direction through the prechamber.
  • a combustion apparatus adapted for use in a gas turbine engine characterized by substantially complete combustion of liquid hydrocarbon fuel and by a low output of nitrogen oxides, the apparatus comprising a combustion liner having a discharge outlet for combustion products at the downstream end of the liner; the liner having an upstream end and liner wall means extending from the upstream end to the downstream end, the wall means enclosing, in sequence from the upstream end to the downstream end, a prechamber, a reaction zone, and a dilution zone, the prechamber including air entrance means at its upstream end effective to direct combustion air with substantial velocity downstream over the inner surface of the prechamber wall means, and liquid fuel introduction means downstream of the said air entrance means disposed to lay a film of liquid fuel on the said inner surface for evaporation by and mixture in the prechamber with the said combustion air, wherein the improvement comprises a textured configuration of the said prechamber wall means inner surface defined by a grid of two sets of intersecting grooves in the surface and generally rectangular bosses rising between adjacent grooves

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Combustion Methods Of Internal-Combustion Engines (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US439648A 1974-02-04 1974-02-04 Combustion apparatus Expired - Lifetime US3859787A (en)

Priority Applications (8)

Application Number Priority Date Filing Date Title
US439648A US3859787A (en) 1974-02-04 1974-02-04 Combustion apparatus
CA212,069A CA1013153A (en) 1974-02-04 1974-10-23 Combustion apparatus for a gas turbine engine
GB5374074A GB1467499A (en) 1974-02-04 1974-12-12 Combustion apparatus
DE2460709A DE2460709C2 (de) 1974-02-04 1974-12-18 Brennkammer für Gasturbinen
IT47881/75A IT1026493B (it) 1974-02-04 1975-01-28 Combustore adatto in pardicolare per motori a turbina a basso livello di inruinamento
SE7500960A SE393839B (sv) 1974-02-04 1975-01-29 Forbrenningsanordning
JP50014046A JPS5825931B2 (ja) 1974-02-04 1975-02-04 ネンシヨウソウチ
FR7503466A FR2259989B1 (enrdf_load_stackoverflow) 1974-02-04 1975-02-04

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JP (1) JPS5825931B2 (enrdf_load_stackoverflow)
CA (1) CA1013153A (enrdf_load_stackoverflow)
DE (1) DE2460709C2 (enrdf_load_stackoverflow)
FR (1) FR2259989B1 (enrdf_load_stackoverflow)
GB (1) GB1467499A (enrdf_load_stackoverflow)
IT (1) IT1026493B (enrdf_load_stackoverflow)
SE (1) SE393839B (enrdf_load_stackoverflow)

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US3925002A (en) * 1974-11-11 1975-12-09 Gen Motors Corp Air preheating combustion apparatus
US3930368A (en) * 1974-12-12 1976-01-06 General Motors Corporation Combustion liner air valve
US3938324A (en) * 1974-12-12 1976-02-17 General Motors Corporation Premix combustor with flow constricting baffle between combustion and dilution zones
US3958416A (en) * 1974-12-12 1976-05-25 General Motors Corporation Combustion apparatus
US3973395A (en) * 1974-12-18 1976-08-10 United Technologies Corporation Low emission combustion chamber
US4004414A (en) * 1973-12-04 1977-01-25 The Franch State Combustion chamber for supercharged internal combustion engine
US4045956A (en) * 1974-12-18 1977-09-06 United Technologies Corporation Low emission combustion chamber
US4058977A (en) * 1974-12-18 1977-11-22 United Technologies Corporation Low emission combustion chamber
US4187674A (en) * 1977-01-21 1980-02-12 Rolls-Royce Limited Combustion equipment for gas turbine engines
US4192139A (en) * 1976-07-02 1980-03-11 Volkswagenwerk Aktiengesellschaft Combustion chamber for gas turbines
FR2452598A1 (fr) * 1979-03-27 1980-10-24 Gen Electric Appareil de combustion comprenant une chambre de precombustion a cyclone, et procede de combustion pour turbines a gaz a combustible liquide
EP0026594A1 (en) * 1979-09-28 1981-04-08 General Motors Corporation Low emissions prevaporization type combustor assembly
US4269466A (en) * 1979-11-23 1981-05-26 Amp Incorporated Connector and strain relief for flat transmission cable
FR2504195A1 (fr) * 1981-04-17 1982-10-22 Gen Electric Systeme d'alimentation en combustible des turbines a gaz
EP0074196A1 (en) * 1981-09-04 1983-03-16 General Motors Corporation Gas turbine prechamber and fuel manifold structure
US4735044A (en) * 1980-11-25 1988-04-05 General Electric Company Dual fuel path stem for a gas turbine engine
US5069029A (en) * 1987-03-05 1991-12-03 Hitachi, Ltd. Gas turbine combustor and combustion method therefor
US5081843A (en) * 1987-04-03 1992-01-21 Hitachi, Ltd. Combustor for a gas turbine
US5669218A (en) * 1995-05-31 1997-09-23 Dresser-Rand Company Premix fuel nozzle
US5794449A (en) * 1995-06-05 1998-08-18 Allison Engine Company, Inc. Dry low emission combustor for gas turbine engines
EP1063476A1 (en) 1999-06-22 2000-12-27 ABB Alstom Power UK Ltd. Combustor for gas turbine engine
US6539721B2 (en) * 2001-07-10 2003-04-01 Pratt & Whitney Canada Corp. Gas-liquid premixer
US20050144929A1 (en) * 2001-11-20 2005-07-07 Volvo Aero Corporation Device for a combustion chamber of a gas turbine
US20070157617A1 (en) * 2005-12-22 2007-07-12 Von Der Bank Ralf S Lean premix burner with circumferential atomizer lip
US20080083223A1 (en) * 2006-10-04 2008-04-10 Lev Alexander Prociw Multi-channel fuel manifold
US20080104961A1 (en) * 2006-11-08 2008-05-08 Ronald Scott Bunker Method and apparatus for enhanced mixing in premixing devices
EP1998035A3 (en) * 2007-05-31 2012-02-08 United Technologies Corporation Fluidic vectoring for exhaust nozzle
CN101514819B (zh) * 2008-02-20 2013-05-15 富来科斯能能源系统公司 气冷旋流式喷嘴头
US8887390B2 (en) 2008-08-15 2014-11-18 Dresser-Rand Company Method for correcting downstream deflection in gas turbine nozzles
US9181813B2 (en) 2012-07-05 2015-11-10 Siemens Aktiengesellschaft Air regulation for film cooling and emission control of combustion gas structure
US20180217070A1 (en) * 2015-07-21 2018-08-02 Fluidsens International Inc. Particles in liquid detection method and particles in liquid detection system and method to detect particles in the air
CN112483260A (zh) * 2020-12-15 2021-03-12 通化师范学院 一种燃气轮机的启动设备
US10989413B2 (en) * 2019-07-17 2021-04-27 General Electric Company Axial retention assembly for combustor components of a gas turbine engine
US10995699B2 (en) * 2018-02-19 2021-05-04 Mra Systems, Llc. Thrust reverser cascade
US11428411B1 (en) 2021-05-18 2022-08-30 General Electric Company Swirler with rifled venturi for dynamics mitigation
US11592178B2 (en) * 2018-05-15 2023-02-28 Air Products And Chemicals, Inc. System and method of improving combustion stability in a gas turbine
US20240263786A1 (en) * 2023-02-02 2024-08-08 Pratt & Whitney Canada Corp. Central air passage with radial fuel distributor

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DE2949388A1 (de) * 1979-12-07 1981-06-11 Kraftwerk Union AG, 4330 Mülheim Brennkammer fuer gasturbinen und verfahren zum betrieb der brennkammer
FR2471480A1 (fr) * 1979-12-13 1981-06-19 Snecma Dispositif d'injection pour chambre a combustion de moteur a turbine
EP0035869B1 (en) * 1980-03-05 1984-07-11 Hitachi, Ltd. A gas turbine combustor
DE3209135A1 (de) * 1982-03-12 1983-09-15 Kraftwerk Union AG, 4330 Mülheim Gasturbinenbrennkammer
JP2516822Y2 (ja) * 1988-08-04 1996-11-13 川崎重工業株式会社 ガスタービン用燃焼器
DE4435266A1 (de) * 1994-10-01 1996-04-04 Abb Management Ag Brenner
US7028484B2 (en) 2002-08-30 2006-04-18 Pratt & Whitney Canada Corp. Nested channel ducts for nozzle construction and the like
US7654088B2 (en) * 2004-02-27 2010-02-02 Pratt & Whitney Canada Corp. Dual conduit fuel manifold for gas turbine engine
US20060156733A1 (en) 2005-01-14 2006-07-20 Pratt & Whitney Canada Corp. Integral heater for fuel conveying member
US7565807B2 (en) 2005-01-18 2009-07-28 Pratt & Whitney Canada Corp. Heat shield for a fuel manifold and method
US7533531B2 (en) 2005-04-01 2009-05-19 Pratt & Whitney Canada Corp. Internal fuel manifold with airblast nozzles
US7540157B2 (en) 2005-06-14 2009-06-02 Pratt & Whitney Canada Corp. Internally mounted fuel manifold with support pins
US7559201B2 (en) 2005-09-08 2009-07-14 Pratt & Whitney Canada Corp. Redundant fuel manifold sealing arrangement
US7854120B2 (en) 2006-03-03 2010-12-21 Pratt & Whitney Canada Corp. Fuel manifold with reduced losses
US7942002B2 (en) 2006-03-03 2011-05-17 Pratt & Whitney Canada Corp. Fuel conveying member with side-brazed sealing members
US7607226B2 (en) 2006-03-03 2009-10-27 Pratt & Whitney Canada Corp. Internal fuel manifold with turned channel having a variable cross-sectional area
US7624577B2 (en) 2006-03-31 2009-12-01 Pratt & Whitney Canada Corp. Gas turbine engine combustor with improved cooling
US8096130B2 (en) 2006-07-20 2012-01-17 Pratt & Whitney Canada Corp. Fuel conveying member for a gas turbine engine
US8353166B2 (en) 2006-08-18 2013-01-15 Pratt & Whitney Canada Corp. Gas turbine combustor and fuel manifold mounting arrangement
US7765808B2 (en) 2006-08-22 2010-08-03 Pratt & Whitney Canada Corp. Optimized internal manifold heat shield attachment
US8033113B2 (en) 2006-08-31 2011-10-11 Pratt & Whitney Canada Corp. Fuel injection system for a gas turbine engine
US7703289B2 (en) 2006-09-18 2010-04-27 Pratt & Whitney Canada Corp. Internal fuel manifold having temperature reduction feature
US7775047B2 (en) 2006-09-22 2010-08-17 Pratt & Whitney Canada Corp. Heat shield with stress relieving feature
US7926286B2 (en) 2006-09-26 2011-04-19 Pratt & Whitney Canada Corp. Heat shield for a fuel manifold
US8572976B2 (en) 2006-10-04 2013-11-05 Pratt & Whitney Canada Corp. Reduced stress internal manifold heat shield attachment
US7856825B2 (en) 2007-05-16 2010-12-28 Pratt & Whitney Canada Corp. Redundant mounting system for an internal fuel manifold
US8146365B2 (en) 2007-06-14 2012-04-03 Pratt & Whitney Canada Corp. Fuel nozzle providing shaped fuel spray
EP3015772B1 (en) * 2014-10-31 2020-01-08 Ansaldo Energia Switzerland AG Combustor arrangement for a gas turbine

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Cited By (46)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US4004414A (en) * 1973-12-04 1977-01-25 The Franch State Combustion chamber for supercharged internal combustion engine
US3925002A (en) * 1974-11-11 1975-12-09 Gen Motors Corp Air preheating combustion apparatus
US3930368A (en) * 1974-12-12 1976-01-06 General Motors Corporation Combustion liner air valve
US3938324A (en) * 1974-12-12 1976-02-17 General Motors Corporation Premix combustor with flow constricting baffle between combustion and dilution zones
US3958416A (en) * 1974-12-12 1976-05-25 General Motors Corporation Combustion apparatus
US3973395A (en) * 1974-12-18 1976-08-10 United Technologies Corporation Low emission combustion chamber
US4045956A (en) * 1974-12-18 1977-09-06 United Technologies Corporation Low emission combustion chamber
US4058977A (en) * 1974-12-18 1977-11-22 United Technologies Corporation Low emission combustion chamber
US4192139A (en) * 1976-07-02 1980-03-11 Volkswagenwerk Aktiengesellschaft Combustion chamber for gas turbines
US4187674A (en) * 1977-01-21 1980-02-12 Rolls-Royce Limited Combustion equipment for gas turbine engines
FR2452598A1 (fr) * 1979-03-27 1980-10-24 Gen Electric Appareil de combustion comprenant une chambre de precombustion a cyclone, et procede de combustion pour turbines a gaz a combustible liquide
EP0026594A1 (en) * 1979-09-28 1981-04-08 General Motors Corporation Low emissions prevaporization type combustor assembly
US4263780A (en) * 1979-09-28 1981-04-28 General Motors Corporation Lean prechamber outflow combustor with sets of primary air entrances
US4269466A (en) * 1979-11-23 1981-05-26 Amp Incorporated Connector and strain relief for flat transmission cable
US4735044A (en) * 1980-11-25 1988-04-05 General Electric Company Dual fuel path stem for a gas turbine engine
FR2504195A1 (fr) * 1981-04-17 1982-10-22 Gen Electric Systeme d'alimentation en combustible des turbines a gaz
EP0074196A1 (en) * 1981-09-04 1983-03-16 General Motors Corporation Gas turbine prechamber and fuel manifold structure
US5069029A (en) * 1987-03-05 1991-12-03 Hitachi, Ltd. Gas turbine combustor and combustion method therefor
US5081843A (en) * 1987-04-03 1992-01-21 Hitachi, Ltd. Combustor for a gas turbine
US5816041A (en) * 1995-05-31 1998-10-06 Dresser Industries, Inc. Premix fuel nozzle
US5669218A (en) * 1995-05-31 1997-09-23 Dresser-Rand Company Premix fuel nozzle
US5813232A (en) * 1995-06-05 1998-09-29 Allison Engine Company, Inc. Dry low emission combustor for gas turbine engines
US5794449A (en) * 1995-06-05 1998-08-18 Allison Engine Company, Inc. Dry low emission combustor for gas turbine engines
EP1063476A1 (en) 1999-06-22 2000-12-27 ABB Alstom Power UK Ltd. Combustor for gas turbine engine
US6425240B1 (en) 1999-06-22 2002-07-30 Abb Alstom Power Uk Ltd. Combustor for gas turbine engine
US6539721B2 (en) * 2001-07-10 2003-04-01 Pratt & Whitney Canada Corp. Gas-liquid premixer
US20050144929A1 (en) * 2001-11-20 2005-07-07 Volvo Aero Corporation Device for a combustion chamber of a gas turbine
US7096675B2 (en) * 2001-11-20 2006-08-29 Volvo Aero Corporation Device for a combustion chamber in a gas turbine for controlling the intake of gas to a combustion zone
US7658075B2 (en) 2005-12-22 2010-02-09 Rolls-Royce Deutschland Ltd & Co Kg Lean premix burner with circumferential atomizer lip
US20070157617A1 (en) * 2005-12-22 2007-07-12 Von Der Bank Ralf S Lean premix burner with circumferential atomizer lip
US7716933B2 (en) 2006-10-04 2010-05-18 Pratt & Whitney Canada Corp. Multi-channel fuel manifold
US20080083223A1 (en) * 2006-10-04 2008-04-10 Lev Alexander Prociw Multi-channel fuel manifold
US20080104961A1 (en) * 2006-11-08 2008-05-08 Ronald Scott Bunker Method and apparatus for enhanced mixing in premixing devices
EP1998035A3 (en) * 2007-05-31 2012-02-08 United Technologies Corporation Fluidic vectoring for exhaust nozzle
CN101514819B (zh) * 2008-02-20 2013-05-15 富来科斯能能源系统公司 气冷旋流式喷嘴头
US9669495B2 (en) 2008-08-15 2017-06-06 Dresser-Rand Company Apparatus for refurbishing a gas turbine nozzle
US8887390B2 (en) 2008-08-15 2014-11-18 Dresser-Rand Company Method for correcting downstream deflection in gas turbine nozzles
US9181813B2 (en) 2012-07-05 2015-11-10 Siemens Aktiengesellschaft Air regulation for film cooling and emission control of combustion gas structure
US20180217070A1 (en) * 2015-07-21 2018-08-02 Fluidsens International Inc. Particles in liquid detection method and particles in liquid detection system and method to detect particles in the air
US11119049B2 (en) * 2015-07-21 2021-09-14 Fluidsens International Inc. Particles in liquid detection method and particles in liquid detection system and method to detect particles in the air
US10995699B2 (en) * 2018-02-19 2021-05-04 Mra Systems, Llc. Thrust reverser cascade
US11592178B2 (en) * 2018-05-15 2023-02-28 Air Products And Chemicals, Inc. System and method of improving combustion stability in a gas turbine
US10989413B2 (en) * 2019-07-17 2021-04-27 General Electric Company Axial retention assembly for combustor components of a gas turbine engine
CN112483260A (zh) * 2020-12-15 2021-03-12 通化师范学院 一种燃气轮机的启动设备
US11428411B1 (en) 2021-05-18 2022-08-30 General Electric Company Swirler with rifled venturi for dynamics mitigation
US20240263786A1 (en) * 2023-02-02 2024-08-08 Pratt & Whitney Canada Corp. Central air passage with radial fuel distributor

Also Published As

Publication number Publication date
GB1467499A (en) 1977-03-16
FR2259989A1 (enrdf_load_stackoverflow) 1975-08-29
CA1013153A (en) 1977-07-05
SE393839B (sv) 1977-05-23
JPS50140716A (enrdf_load_stackoverflow) 1975-11-12
IT1026493B (it) 1978-09-20
FR2259989B1 (enrdf_load_stackoverflow) 1980-01-11
DE2460709A1 (de) 1975-08-07
SE7500960L (enrdf_load_stackoverflow) 1975-08-05
JPS5825931B2 (ja) 1983-05-31
DE2460709C2 (de) 1982-12-16

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