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Publication numberUS3851465 A
Publication typeGrant
Publication dateDec 3, 1974
Filing dateApr 6, 1973
Priority dateApr 6, 1973
Also published asCA982830A1
Publication numberUS 3851465 A, US 3851465A, US-A-3851465, US3851465 A, US3851465A
InventorsA Verdouw
Original AssigneeGen Motors Corp
Export CitationBiBTeX, EndNote, RefMan
External Links: USPTO, USPTO Assignment, Espacenet
Annular dilution zone combustor
US 3851465 A
Abstract  available in
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Claims  available in
Description  (OCR text may contain errors)

Umted States Patent 1191 1111 3,851,465 Verdouw Dec. 3, 1974 4] ANNULAR DILUTION ZONE COMBUSTOR 3,169,369 2/1965 Holl 60/3966 3,518,037 6/1970 Sneeden..... [751 lmemmi i 3,608,309 9/1971 111116161. 60/39.65

[73] Assignee: General Motors Corporation, m y Examiner-Carlton R- Croyle Detroit, Mich. Assistant Examiner-Robert E. Garrett Filed: p 1973 Attorney, Agent, or FzrmPaul Fltzpatrick [21] App]. No.: 348,682 [57] ABSTRACT A combustion liner of the single can type for dis- 521 US. (:1 60/39.36, 60/39.65, 60/39.66, charge into an annular turbine nozzle is of the same 60/DIG 11 overall dimensions as a prior art liner for the same in- 51 Int. Cl F020 3/04,1=02 7/18 stell'aheh, but differs in the elimination of dilution air 58 Field 61 Search 60/DIG. 11, 39.65, 39.66, holes hear the middle of the length of the liner and 0 9 provision of a ring of dilution air holes near the outlet 1 end of the liner. Also, a centerbody mounted in the 5 References Cited outlet end of the liner provides an annular dilution UNITED STATES PATENTS zone into which the dilution air flowing through the holes is discharged. 2,765,621 10/1956 Poulston et al. 60/3965 X 2,781,638 2/1957 Fletcher et al 60/39.65 x 5 Claims, 4 Drawing g s 1 Z [Z l/ 4" 11 7 5 i 61% A; I! Z 6 42 Z' i Z2 1 dz; 24 i 32, a 2.9

I a l 1 3 I at if I z 5 w I t "I i Z07 1 1 I l I Z2 ANNULAR DILUTION ZONE COMBUSTOR The invention disclosed and claimed herein was made in the course of work under a contract with the Department of Defense.

My invention is directed to improvements in combustion apparatus, particularly combustion apparatus of such characteristics as are required in gas turbine engines. Most particularly, the invention is directed to combustion apparatus of a type in which a single combustion liner discharges through the entire circumference of an annular turbine entrance.

One application of such combustion liners is in the well-known T63 turboshaft engine, widely applied in light helicopters. This invention was conceived in the course of work to improve the quality of the exhaust of such an engine; that is, to reduce the output of substances regarded as atmospheric pollutants. The particular-object of the work leading to this invention was to improve exhaust quality without changes in the dimensions of the combustion liner which would require enlargement of the engine combustion apparatus and thus add to the overall length of the power plant.

The conventional approach in designing gas turbine combustion apparatus has been to secure efficient combustion; that is, substantially complete combustion while minimizing pressure drop, and while keeping the combustion apparatus as short as possible to minimize weight and improve installation characteristics of the gas turbine engine. This general consideration applies to engines of many types. Recent emphasis on reduction or elimination of substances such as particulates, unburned hydrocarbons, carbon monoxide, and oxides of nitrogen have led to reconsideration of the principles of combustor design.

It has long been realized that a long combustion apparatus may readily be made to burn more cleanly than a short one, and thus some approaches to improving the emission characteristics of gas turbine combustion chambers have led to quite substantial increases in combustion liner and combustion apparatus length. Increase in length of the combustion apparatus makes possible a longer combustion zone with greater residence time of the gases in the combustion zone, resulting in more complete combustion of the fuel.

Approaching this problem with a view to improving the exhaust quality while retaining the advantages of compact combustion apparatus and short overall engine length, I have conceived of a combustion liner structure which provides a much increased length combustion zone in the same overall length combustion liner while maintaining effective dilution of the combustion products. It is, of course, necessary to dilute the I combustion products before they are admitted to the turbine to reduce the temperature of the motive fluid. It is unfeasible to supply all the air initially to the combustion zone, as this results in too low ,a fuel-air ratio for proper combustion; therefore standard combustors burn the fuel in part of the air (primary air) and then add dilution air down-stream.

The principal objects of my invention are to improve the quality of exhaust of gas turbine engines, to provide a combustor having lower outputs of unburned hydrocarbons and carbon monoxide than conventional combustors, to provide a combustion liner of improved combustion characteristics having the same overall dimensions as a conventional liner for which it may be substituted, to provide a combustion apparatus of short overall length with relatively great length of the combustion zone and short length of the dilution zone, to provide a combustion apparatus having a liner with a circular cross section combustion zone and an annular cross section dilution zone, and in general to provide a combustion apparatus better suited to the requirements of practice than those presently available.

The nature of my invention and its advantages will be more clearly apparent to those skilled in the art from I the succeeding detailed description of the preferred embodiment and the accompanying drawings;

FIG. 1 is a longitudinal sectional view of a combustion apparatus of a gas turbine engine.

FIG. 2 is a partial exterior view of a combustion line r taken on the plane indicated by the line 22 in FIG. 1.

FIG. 3 is an end elevation view of the liner taken on the plane indicated by the line 33 in FIG. 2.

FIG. 4 is a cross sectional view of a portion of the liner taken on the plane indicated by the line 4-4 in FIG. 1.

Referring first to FIG. 1, this shows a combustion liner according to the invention as installed in a T63 engine. The engine includes an outer case 2 to which is bolted the casing 3 of a combustion apparatusQCombustion takes place within a liner 4 mounted concentrically within the generally cylindrical casing 3, and the combustion products are discharged through a turbine inlet 6 defined by outer and inner walls 7 and 8. The inner wall 8 is joined at its upstream end to a baffle 10 disposed upstream of the turbine rotor disks and bearings.

In this engine, compressed air is supplied through air tubes (not illustrated) into an entrance zone 11 of the combustion casing, from which it flows through a perforated baffle 12 into proximity to the liner 4.'The liner includes a conical dome 14 at its upstream end, a first wall section 15, and a second wall section 16. The wall sections 15 and 16 constitute the outer wall of the liner which is approximately cylindrical but is slightly convergent-divergent as shown. The forward ends of the sections 15 and 16 arecorrugated and are spot-welded to the downstream end of section 15 and the periphery of the dome respectively, to provide film cooling air entrances 18 as described in Hayes US. Pat. No. 3,064,425 issued Nov. 20, 1962. A ferrule 19 at the center of the dome is piloted on the discharge end of a fuel spray nozzle 20 mounted in a boss at the upstream end of casing 3. Primary or combustion air is admitted to the liner through perforations 22 distributed over the surface of the dome and through a ring of twelve primary air holes 23 in wall section 15. An igniter (not illustrated) may be installed in any usual manner.

. The downstream end of liner 4 pilots within the turbine outer wall 7. A sleeve 24 welded to the wall section 16 fits over the turbine wall 7. It may be pointed out that the liner illustrated, as described so far, is identical to the prior art combustion liner. The overall length of liner 4 is 9 /2 inches. Further to provide background as to prior art, it may be pointed out that the prior art T63 combustion liner has a ring of 14 trim air holes three-eight inch in diameter at the station along the liner indicated by the dotted line 26. The prior liner has also two opposed 1%. inch diameter dilution air holes at the station indicated at 27. In the present liner this portion is imperforate. Approximately, the air admitted through the holes at locations 26 and 27 in the prior art liner is admitted as dilution air through a ring of dilution holes or slots 28 (see also FIG. 2) located as shown near the downstream end of my liner. Specifically, these slots are 0.71 inch X 0.28 inch in size.

My liner also includes a cup-shaped centerbody 30 mounted upstream of the baffle and radially inwardly of the dilution holes 28. The centerbody includes main body 31, a downstream closure wall 32, and cooling air baffles 34 and 35 on the upstream face of the centerbody.

The centerbody is supported by four turbular struts 36 open at both ends which are welded or brazed to the liner wall section 16 and the side wall of the center body. In addition to supporting the centerbody, the struts conduct cooling air from the interior of easing 3. The cooling air is discharged through 8 holes 38 behind baffle 35 and 36 holes 39 behind baffle 34. These holes are one-eighth inch in diameter. Baffle 35 is fixed to a stud 40 extending from the face of the centerbody and baffle 34 is welded to the face of the centerbody. These baffles deflect the cooling air radially outward over the face of the centerbody and the air also flows as a film over the outer surface of body 31 toward the outlet of the combustor. Cooling of the centerbody is needed because it is exposed to very intense heat radiation from the combustion zone upstream of the centerbody and convection from the hot gases flowing over it.

It is contemplated that the centerbody may be porous, in which case it may be cooled by transpiration of air from the interior of the centerbody through closely spaced pores distributed over the surface of the centerbody. The centerbody for transpiration cooling might be made of porous ceramic or of porous laminated material of the type described in U.S. Pat. 3,584,972 of Bratkovich and Meginnis issued June 15, 1971.

With the structure shown the flame is not quenched by the dilustion air until it reaches the zone of the dilution air holes 28.'This is a quite significantly longer combustion zone than in the prior art combustor in which combustion was quenched in the region indicated at 26 and 27 in FIG. 1, as previously described.

The centerbody, by narrowing the passage toward the outlet of the combustion chamber provides a much higher length to diameter ratio for the passage in which the dilution air is mixed, thus contributing to effective mixing in a relatively short distance from holes 28 into the outlet of the combustion liner. Since an effective dilution zone is shorter, a longer combustion zone is available in the same overall liner length. 1 consider that the combustor according to the invention increases the reaction length for consuming carbon monoxide, hydrocarbons, and carbon over the prior liner from llinches to 4% inches prior to the final quench.

The primay zone equivalence ratio (actual fuel-air ratio divided by stoichiometric ratio) of the prior art liner was 0.77 and that of the liner described here 0.80 at maximum power conditions. The distribution of air flow for the prior liner ad for my improved liner are as tabulated below Location Prior Art Improved Liner Dome holes 11.8% 11.3% First cooling step 11.2% 10.771 Primary holes 26.3% 25.4% Second cooling step 11.271 10.771 Trim holes 15.271 Dilution holes 24.371 38.07: centerbody cooling 3.971

Total 100.071 100.071

Tests have been made of my combustor with regard to a stipulated duty cycle (relative operating times at diverse power levels) of a T63 engine in an aircraft. These tests demonstrated significant reductions in carbon monoxide and hydrocarbons with substantially the same oxides of nitrogen content and somewhat higher smoke level. The combustor with the improved liner reduced total emissions relative to the prior liner by one-third, which is a significant improvement. While a greater degree of reduction of emissions can be had with other forms of combustor, it should be borne in vmind that this improvement is accomplished without any change in dimensions of the combustion apparatus or any modification of the engine other than to the combustion liner itself. It will be seen, therefore, that this invention can result in a substantial improvement in the engine.

It will be appreciated that the specific dimensions and proportions stated above are illustrative, and may be varied to suit particular installations.

U.S. Pat. No. 3,169,369 is directed to a single-can combustion apparatus with annular outlet into a turbine. It will be apparent, however, that it is quite remote from my combustion apparatus.

The detailed description of the preferred embodiment of the invention for the purpose of explaining the principles thereof is not to be considered as limiting the invention, as many modifications may be made by the exercise of those skilled in the art.

1 claim:

1. A combustion liner for a gas turbine combustion apparatus, the liner including an outer wall of generally circular cross-section having an upstream end and a downstream end, a dome closing the upstream end, and a substantially closed centerbody mounted in the downstream end, the centerbody and outer wall defining between them an annular outlet passage of substantially constant width from the liner; the dome providing for entrance of fuel; approximately the upstream third of the liner including the dome providing distributed entrances for combustion air and cooling air; at least approximately the central third of the liner wall being imperforate to avoid quenching of the combustion; and the outer wall having a ring of dilution air holes abreast of the centerbody to admit dilution air directly into the outlet passage for mixing with and cooling of the combustion products flowing through the outlet passage.

2. A combustion liner for a gas turbine combustion apparatus, the liner including an outer wall of substantially circular cross-section having an upstream end and a downstream end, a dome closing the upstream end, and a substantially closed centerbody mounted in the downstream end, the centerbody and outer wall defining between them an annular outlet passage from the liner; the dome providing for entrance of fuel; approximately the upstream third of the liner including the dome providing distributed entrances for combustion air and cooling air; at least approximately the central third of the liner wall being imperforate to avoid quenching of the combustion; the said outlet passage having a radial width less than its length axially of the liner; and the outer wall having a ring of dilution air holes abreast of the centerbody to admit dilution air directly into the outlet passage for mixing with and cooling of the combustion products flowing through the outlet passage.

3. A combustion liner for a gas turbine combustion apparatus including an approximately cylindrical wall, a dome defining the upstream end of the liner and providing a seat for fuel spray means, and a centerbody extending into the downstream end of the liner coaxially with the liner wall; the wall and centerbody defining between them an annular passage of substantially constant width substantially less than the radius of the liner wall leading to an annular outlet from the liner; the space defined by the liner upstream of the centerbody being a combustion space having inlets for primary combustion air, the liner wall having a ring of dilution air openings abreast of the centerbody for entrance of dilution air directly into the said annular passage and mixing of the dilution air therein with the combustion products flowing from'the combustion space.

4. A combustion apparatus comprising, in combination, an outer case adapted to receive compressed air and a combustion liner within the case, the liner including an approximately cylindrical wall, a dome defining the upstream end of the liner and providing a seat for 6 fuel spray means, and a centerbody extending into the downstream end of the liner coaxially with the liner wall; the wall and centerbody defining between them an annular passage, of width less than half the radius of the liner wall, leading to an annular outlet from'the liner; the space defined by the liner upstream of the centerbody being a combustion space having inlets for primary combustion air; the liner wall having a ring of dilution air openings abreast of the centerbody for entrance of dilution air directly into the said annular passage and mixing of the dilution air therein with the combustion products flowing from the combustio space. 5. A combustion apparatus comprising, in combination, an outer case adapted to receive compressed air and a combustion liner within the case, the liner including an approximately cylindrical wall, a dome defining the upstream end of the liner and providing a seat for fuel spray means, and a centerbody extending into the downstream end of the liner coaxially with the liner wall; the wall and centerbody defining between them an annular passage, of substantially constant width less than half the radius of the liner wall and of a length of the order of one-fourth the length of the liner, leading to an annular outlet from the liner; the space defined by the liner upstream of the centerbody being a combustion space having inlets for primary combustion air; the liner wall having a ring of dilution air openings abreast of the centerbody for entrance of dilution air directly into the said annular passage and mixing of the dilution air therein with the combustion products flowing from the combustion space.

Referenced by
Citing PatentFiling datePublication dateApplicantTitle
US4171612 *Jun 6, 1977Oct 23, 1979Zwick Eugene BLow emission burner construction
US4173118 *Aug 31, 1977Nov 6, 1979Mitsubishi Jukogyo Kabushiki KaishaFuel combustion apparatus employing staged combustion
US4292810 *Feb 1, 1979Oct 6, 1981Westinghouse Electric Corp.Gas turbine combustion chamber
US4903492 *Sep 7, 1988Feb 27, 1990The United States Of America As Represented By The Secretary Of The Air ForceDilution air dispensing apparatus
US5220795 *Apr 16, 1991Jun 22, 1993General Electric CompanyMethod and apparatus for injecting dilution air
US5277021 *May 13, 1991Jan 11, 1994Sundstrand CorporationVery high altitude turbine combustor
US5373694 *Aug 9, 1993Dec 20, 1994United Technologies CorporationCombustor seal and support
US5456080 *Jan 7, 1994Oct 10, 1995Sundstrand CorporationVery high altitude turbine combustor
US8397511 *May 19, 2009Mar 19, 2013General Electric CompanySystem and method for cooling a wall of a gas turbine combustor
US20100293957 *May 19, 2009Nov 25, 2010General Electric CompanySystem and method for cooling a wall of a gas turbine combustor
EP1843097A1 *Mar 3, 2007Oct 10, 2007Siemens Power Generation, Inc.Air flow conditioner for a combustor can of a gas turbine engine
Classifications
U.S. Classification60/804, 60/755
International ClassificationF23R3/54, F23R3/42, F23R3/04
Cooperative ClassificationF23R3/42, F23R3/04, F23R3/54, Y02T50/675
European ClassificationF23R3/04, F23R3/42, F23R3/54