US20040011058A1

US20040011058A1 - Annular combustion chamber with two offset heads

Info

Publication number: US20040011058A1
Application number: US10/227,815
Authority: US
Inventors: Christophe Baudoin; Patrice Commaret; Eric Le Letty; Christophe Viguier
Original assignee: SNECMA Moteurs SA
Current assignee: Safran Aircraft Engines SAS
Priority date: 2001-08-28
Filing date: 2002-08-27
Publication date: 2004-01-22
Also published as: EP1288579A1; JP4038094B2; JP2003106529A; FR2829228A1; UA80669C2; EP1288579B1; ES2300426T3; FR2829228B1; DE60225095D1; CA2398669A1; CN1407279A; RU2296917C2; UA79922C2; DE60225095T2; RU2002123305A; CA2398669C

Abstract

An offset annular combustion chamber for an airplane engine gas turbine, the chamber comprising a pilot head having a plurality of nozzle systems distributed on a pilot head chamber end wall interconnecting an inner longitudinally-extending side wall of the chamber to a pilot head outer longitudinally-extending side wall, and a take-off head that is radially and axially offset from the pilot head and comprising a plurality of nozzle systems distributed on a take-off head chamber end wall interconnecting the pilot head outer longitudinally-extending side wall and a take-off head outer longitudinally-extending side wall, the pilot head having at least N substantially identical nozzle systems with overall permeability PA adapted to lighting and to speeds close to idling, and the take-off head having at least 2N substantially identical nozzle systems of overall permeability PB, where PB is greater than or equal to PA, the permeability PA lying in the range 10% to 40% of the total air flow rate penetrating into the chamber and the permeability PB lying in the range 30% to 70% of the total air flow rate penetrating into the chamber.

Description

BACKGROUND OF THE INVENTION

The present invention relates to the field of annular combustion chambers with two offset heads for an airplane engine gas turbine. The invention relates more particularly to the arrangement of fuel nozzle systems fitted to such combustion chambers.

A gas turbine combustion chamber is formed in conventional manner by inner and outer annular side walls extending longitudinally and united by a chamber end wall. Combustion reactions take place within the chamber, and the chamber is configured so that the flow of air that it receives is shared between at least three fractions: combustion air; dilution air; and air needed for cooling the chamber walls, which air does not participate directly in combustion phenomena. The chamber thus comprises a primary or combustion zone and a secondary zone known as a “dilution” zone which is situated downstream from the preceding zone.

Fuel is fed to the combustion chamber via fuel nozzles placed in holes that pass through the end wall of the chamber. Air for the combustion zone is introduced into the chamber, partly through its end wall and possibly via the nozzles, and partly via through orifices pierced in the longitudinally-extending side walls. Dilution air is generally introduced further downstream in the combustion chamber via one or more rows of holes that are likewise distributed in the side walls of the chamber.

By their very design, the combustion chambers that are presently in use make it difficult to minimize polluting emissions that come from combustion, in particular emissions of nitrogen, of carbon monoxide, and of unburnt hydrocarbons.

In order to solve this problem, it is known to use combustion chambers having two heads that are offset, i.e. chambers in which the fuel nozzles are shared between a so-called “pilot” head and a so-called “take-off” head that is spaced apart from the pilot head both radially and axially, the take-off head being situated downstream from the pilot head in the direction of gas flow in the chamber.

Conventionally, the “pilot” nozzles are used for lighting and when the engine is idling, while the “takeoff” nozzles are used during “full throttle” (FT) stages, in particular during take-off and while cruising. In general, the pilot nozzles are fed with fuel continuously, while the take-off nozzles are fed only above some determined minimum speed.

Thus, for example, document FR 2 727 193 discloses an annular combustion chamber in which the nozzles are distributed over a pilot head and a take-off head. The pilot head is fitted with n nozzles of permeability P1 adapted to idling the speed. The take-off head is also fitted with n nozzles of permeability P1 enabling the take-off head to be lighted at low speed, plus n take-off nozzles of permeability P2>P1 adapted to full load conditions (where the term “permeability” concerning n nozzles relates to the total flow of air passing through all n nozzles).

The permeability P1 of the n pilot nozzles lies in the range 10% to 12% of the total air flow that penetrates into the combustion chamber. Given the head loss due to the air going past the pilot head in order to reach the take-off head, the same permeability P1 for the n first take-off nozzles corresponds to about 8% to 10% of the total incoming air flow. In contrast, the permeability P2 of the n second take-off nozzles is 26% to 35% of said total air flow.

Such a disposition makes it easier to switch between idling conditions, in which only n pilot nozzles are fed, and a sector burning (SB) situation in which only the n take-off nozzles of permeability P1 are lighted amongst the take-off nozzles.

However, because of the large difference between the permeability P2 and the permeability P1 of the take-off nozzles, the subsequent switch from sector burning (SB) to full throttle (FT) is more difficult. It can be achieved only with feed that is relatively rich, and thus with the engine turning at high speed (it is stated in document FR 2 727 193 that the n take-off nozzles of permeability P2 are lighted once the high pressure compressor has reached 70% of its nominal full throttle speed of rotation).

Unfortunately, prolonged operation under sector burning conditions presents drawbacks: temperature distribution over the high pressure turbine blades is not optimum, and alternating lighted nozzles and nozzles that are out in the take-off head encourages chemical reactions to “freeze”, thus affecting combustion efficiency and encouraging undesirable emissions of particles and unburnt fuel.

OBJECT AND SUMMARY OF THE INVENTION

The present invention seeks to mitigate those drawbacks by proposing an annular combustion chamber with two offset heads that provides an operating range that is significantly extended compared with conventional technologies using one or two heads, while nevertheless giving control over temperature profiles and reducing polluting emissions.

To this end, the invention provides an offset annular combustion chamber for an airplane engine gas turbine, the chamber comprising a pilot head having a plurality of nozzle systems distributed on a pilot head chamber end wall interconnecting an inner longitudinally-extending side wall of the chamber to a pilot head outer longitudinally-extending side wall, and a take-off head that is radially and axially offset from the pilot head and comprising a plurality of nozzle systems distributed on a take-off head chamber end wall interconnecting the pilot head outer longitudinally-extending side wall and a take-off head outer longitudinally-extending side wall, the pilot head having at least N substantially identical nozzle systems with overall permeability PA adapted to lighting and to speeds close to idling, and the take-off head having at least 2N substantially identical nozzle systems of overall permeability PB, where PB is greater than or equal to PA, wherein the permeability PA lies in the range 10% to 40% of the total air flow rate penetrating into the chamber and the permeability PB lies in the range 30% to 70% of the total air flow rate penetrating into the chamber.

Using 2N nozzle systems having the same individual permeability for the take-off head makes it possible to ensure that changeover takes place under good conditions both from idling to SB and from SB to FT, where it is possible to perform the SB to FT changeover while running slowly, even close to idling.

In advantageous dispositions, the permeability PA lies in the range 17% to 21% of the total air flow penetrating into the combustion chamber, and the permeability PB lies in the range 36% to 45% of the same air flow.

The outer longitudinally-extending side wall of the take-off head and possibly also the outer longitudinally extending side wall of the pilot head, together with the inner side wall advantageously have rows of dilution orifices. The air flow rate penetrating via said dilution orifices lies in the range 4% to 10% and preferably in the range 6% to 8% of the total air flow rate entering the chamber for the outer orifices in the outer longitudinally-extending side wall(s), and lies in the range 2% to 8% and preferably in the range 4% to 6% for the inner orifices formed in the inner side wall.

The axes of the nozzle systems of the pilot and take-off heads are advantageously directed towards a common annular zone for exhausting the gases that come from combustion.

The nozzle systems of the pilot and take-off heads are installed on end walls of the chamber which may be perpendicular to the axis of the engine or which may be conical in shape.

BRIEF DESCRIPTION OF THE DRAWINGS

Other characteristics and advantages of the present invention appear from the following description given with reference to the accompanying drawings which show an embodiment that is not limiting in any way. In the figures: [0019]
FIG. 1 is a highly diagrammatic axial half-section view of a combustion chamber constituting an embodiment of the invention; [0020]
FIG. 2 is a highly diagrammatic fragmentary view showing one example of how the nozzle systems can be distributed on the end walls of the pilot and take-off heads; and [0021]
FIG. 3 is a section view showing a particular embodiment of a nozzle system in accordance with the invention.[0022]

DETAILED DESCRIPTION OF AN EMBODIMENT

Reference is made initially to FIG. 1 which is a diagrammatic axial half-section showing a [0023] combustion chamber 1 constituting an embodiment of the invention. The axis referenced X-X corresponds to the axis of the engine fitted with such a combustion chamber.
The [0024] combustion chamber 1 is of the annular type with a pilot head 12, and with a take-off head 14 that is offset from the pilot head both radially and axially, the take-off head being situated downstream from the pilot head in the direction of gas flow in the chamber. The chamber is formed in particular by an outer longitudinally-extending side wall 2 for the take-off head, an inner longitudinally-extending side wall 4, and an outer longitudinally-extending side wall 6 for the pilot head. The transverse end wall 8 of the chamber for the pilot head unites the pilot outer side wall 6 and the inner side wall 4, while the take-off head outer side wall 2 and the pilot head outer side wall 6 are united by an end wall 10 of the take-off head, likewise extending transversely.
As shown in FIG. 2, [0025] fuel nozzle systems 16, 18 are placed in holes 16 a and 18 a passing through the respective end walls 8 and 10 of the pilot and take-off heads. More precisely, the pilot head 12 has N fuel nozzle systems 16 that are substantially identical and that are regularly distributed around the axis X-X, while the take-off head 14 has 2N nozzle systems 18 that are substantially identical and that are likewise regularly distributed around the axis X-X.
Thus, in an angular sector of the chamber corresponding to 2π/N, there are to be found two [0026] nozzle systems 18 of the take-off head for each nozzle system 16 of the pilot head. The nozzle systems are advantageously disposed substantially in a staggered configuration. The term “staggered” is used to mean that in an angular sector of 2π/N, the angular position of the pilot head nozzle system 16 is situated at substantially equal distances from the angular positions of the two nozzle systems 18 of the take-off head.
The pilot and take-off [0027] heads 12 and 14 may be fitted with any conventional type of nozzle system serving either to spray fuel in mechanical, aerodynamic, or premixed manner, or else to vaporize fuel. A particular embodiment of a nozzle system is described below with reference to FIG. 3.
The N nozzle systems of the [0028] pilot head 12 have overall permeability PA while the 2N nozzle systems of the take-off head 14 have overall permeability PB, where PB is greater than or equal to PA. Advantageously, 2PA≦PB≦3PA, and preferably 2.5PA≦PB≦3PA.
The term “overall” permeability, PA or PB, is used to designate the air flow rates passing respectively through all N nozzle systems of the pilot head and through all 2N nozzle systems of the take-off head, i.e. PA=Npa, where pa is the individual permeability of each nozzle system of the pilot head, and PB=2Npb where pb is the permeability of each nozzle system of the take-off head. These permeabilities are expressed herein as percentages of the total air flow rate penetrating into the combustion chamber. [0029]
In advantageous dispositions, the permeability PA lies in the [0030] range 10% to 40% and preferably in the range 17% to 21% of the total air flow rate penetrating into the combustion chamber, while the permeability PB lies in the range 30% to 70%, and preferably in the range 36% to 45% of the same air flow rate.
In conventional manner, the longitudinally-extending [0031] side wall 2 of the take-off head and the inner side wall 4 may each be pierced by at least one respective row of dilution orifices of diameter that is adjusted as a function of the required performance. These dilution orifices enable the combustion chamber to be fed with the air required for diluting the combustion gases.
The dilution orifices are preferably distributed as follows: [0032]
the outer longitudinally-extending [0033] side wall 2 of the take-off head has at least one row of 2N outer dilution orifices 20, e.g. identical orifices, opening out into the combustion chamber 1 substantially perpendicularly to the side wall, downstream from the take-off head, and angularly distributed in regular manner about the axis X-X; and
the [0034] inner side wall 4 has at least one row of 2N inner dilution orifices 22 opening out into the combustion chamber substantially perpendicularly to the side wall and angularly distributed in regular manner about the axis X-X.
The [0035] 2N dilution holes 22 may be distributed as N first identical holes occupying the same angular positions as every other nozzle system 18, and N second identical holes occupying the same angular positions as the remaining nozzle systems 18, with the first dilution holes being identical or not identical to the second dilution holes.
The outer longitudinally-extending [0036] side wall 6 of the pilot head may also be provided with outer dilution orifices 20′, and additional inner dilution orifices 22′ could then be provided through the inner wall 4, substantially at the same distance along the combustion chamber, in register with the orifices 20′.
As an indication, the fraction of the air flow rate penetrating into the combustion chamber via the [0037] dilution orifices 20 situated in the outer longitudinally-extending wall 2 of the take-off head (and together with the orifices 20′, if any, situated in the outer longitudinally-extending side wall 6) may lie in the range 4% to 10% and preferably in the range 6% to 8% of the total flow rate, while the flow rate through the orifices situated in the inner side wall 4 may lie in the range 2% to 8%, and preferably in the range 4% to 6% of the same flow rate.
The remaining air flow rate is for cooling the longitudinally-extending side walls and the end walls of the chamber. For this purpose, the longitudinally-extending [0038] side walls 2, 4, and 6 of the chamber are conventionally cooled by these walls being multiply perforated or by fitting these walls with tile devices or with film devices.
In addition, if the chamber is lighted by means of a conventional device (not shown) of the semiconductor or air igniter plug type, it can be placed on the axis of one of the [0039] nozzle systems 16 of the pilot head 12, for example.
Advantageously, the [0040] chamber end walls 8, 10 and the nozzle systems 16, 18 which pass through them are disposed in such a manner that the axes of the nozzle systems point towards a common annular zone for exhausting the gases generated by combustion. For this purpose, FIG. 1 shows two possible dispositions for the chamber end walls and their respective nozzle systems: in continuous lines the end walls 8 and 10 are substantially perpendicular to the axis X-X of the engine, while in dashed lines they are substantially frustoconical in shape. In the first case, the axes Y, Z of the nozzle systems 16, 18 can be inclined relative to the normal to the chamber end walls 8 and 10, whereas in the second case they can be perpendicular to the chamber end walls.
These dispositions serve to reduce as far as possible the total volume of the combustion chamber and to improve performance in terms of temperature, combustion efficiency, and reducing polluting emissions. The convergence of the nozzle system axes serves to increase mixing speed and the rate at which fuel burns in the chamber, and consequently encourages complete combustion of the fuel in a small volume. Since the production of nitrogen oxides is a function of the transit time taken by combustion gases to pass through the combustion chamber, high speed combustion serves to reduce polluting emissions to a significant extent. [0041]
FIG. 3 shows an embodiment of a nozzle system. It comprises a [0042] nozzle 24 that is fed with fuel. Primary and secondary air swirlers 26 and 28 are disposed in such a manner as to feed the nozzle system with air in a radial direction. A Venturi 30 placed on the axis of the fuel nozzle 24 between the primary and secondary swirlers encourages the fuel to break up into a spray of fine droplets. Ventilation holes 32 which open out around and close to the tip of the nozzle 24 serve to limit or even to eliminate any risk of coking on the tip.
The set of nozzle systems on the pilot head may also be provided with fairing typically constituted by two [0043] caps 34 a and 34 b. This fairing serves to minimize head losses in the air that goes round the pilot head and serves to guarantee good air feed to the end wall of the take-off head.
It should be observed that the combustion chamber may be made of ceramic matrix composite (CMC) material which, because of its ability to withstand high temperatures, makes it possible to achieve substantial savings in terms of cooling air flow rate. [0044]
A combustion chamber with two heads that are offset and fitted with nozzle systems in accordance with the invention can operate in the following modes: [0045]
idling mode (or N/0) mode: fuel is injected solely via the pilot head fitted with its N nozzle systems of permeability PA. This mode is intended more particularly for lighting the engine and for operating it at speeds close to idling; [0046]
full throttle mode (FT or N/2N mode): fuel is injected over all of the nozzle systems with it being possible to modulate the way fuel is shared between the pilot and take-off heads. This mode is intended to cover most of the operating range of the chamber and it provides performance that is better in terms of temperature, efficiency, and reducing polluting emissions; and [0047]
sector burning mode (SB or N/N mode): fuel is fed to all of the nozzle systems in the pilot head, and in general, to every other nozzle system in the take-off head. This mode of operation makes it easier to switch between the pilot and take-off heads, particularly when the chamber end walls are of high permeability. [0048]
The disposition of the nozzle systems makes it possible to obtain an operating range for the combustion chamber that is significantly extended, and to obtain lighting performance and stability that are equivalent to or better than a conventional chamber. Furthermore, the transition between SB mode and FT mode can take place at low speed. The take-off nozzles all have the same individual permeability such that switching from SB mode (N/N) to FT mode (N/2N) is easier than in the engine of [0049] document FR 2 727 193 mentioned at the beginning of this description where the n additional take-off nozzles present overall permeability that is much greater than that of the n first nozzles.
The number of nozzle systems in the pilot head (i.e. N) is optimized so as to reconcile lighting and stability performance and flame propagation while still allowing [0050] 2N nozzle systems to be installed in the take-off head. As an indication, the pilot head may be fitted with 16 nozzle systems and the take-off head with 32 nozzle systems.

Claims

1/ An offset annular combustion chamber for an airplane engine gas turbine, the chamber comprising a pilot head having a plurality of nozzle systems distributed on a pilot head chamber end wall interconnecting an inner longitudinally-extending side wall of the chamber to a pilot head outer longitudinally-extending side wall, and a take-off head that is radially and axially offset from the pilot head and comprising a plurality of nozzle systems distributed on a take-off head chamber end wall interconnecting the pilot head outer longitudinally-extending side wall and a take-off head outer longitudinally-extending side wall, the pilot head having at least N substantially identical nozzle systems with overall permeability PA adapted to lighting and to speeds close to idling, and the take-off head having at least 2N substantially identical nozzle systems of overall permeability PB, where PB is greater than or equal to PA, wherein the permeability PA lies in the range 10% to 40% of the total air flow rate penetrating into the chamber and the permeability PB lies in the range 30% to 70% of the total air flow rate penetrating into the chamber.

2/ A chamber according to claim 1, wherein the permeability PA lies in the range 17% to 21% of the total air flow rate penetrating into the chamber.

3/ A chamber according to claim 1, wherein the permeability PB lies in the range 36% to 45% of the total air flow rate penetrating into the chamber.

4/ A chamber according to claim 1, further comprising a plurality of external dilution orifices opening out at least through the outer longitudinally-extending side wall of the take-off head.

5/ A chamber according to claim 4, having at least one row of 2N outer dilution orifices opening out substantially perpendicularly to the outer longitudinally-extending side wall of the take-off head.

6/ A chamber according to claim 4, wherein the air flow rate penetrating via the outer dilution orifices lies in the range 4% to 10% of the total air flow rate penetrating into the chamber.

7/ A chamber according to claim 1, having a plurality of inner dilution orifices opening out through the inner longitudinally-extending side wall of the chamber.

8/ A chamber according to claim 7, wherein the air flow rate penetrating via the inner dilution orifices lies in the range 2% to 8% of the total air flow rate penetrating into the chamber.

9/ A chamber according to claim 1, wherein the nozzle systems of the pilot head and of the take-off head are disposed substantially in a staggered configuration.

10/ A chamber according to claim 1, wherein the axes of the nozzle systems of the pilot head and of the take-off head are directed towards a common annular zone for exhausting the gases generated by combustion.

11/ A chamber according to claim 1, wherein the chamber end walls of the pilot and take-off heads are walls extending perpendicularly to the axis of the engine.

12/ A chamber according to claim 1, wherein the chamber end walls of the pilot and take-off heads are walls of frustoconical shape.

13/ A chamber according to claim 1, wherein each nozzle system of the pilot and take-off heads comprises a fuel nozzle, a primary air swirler and a secondary air swirler that are fed radially, a Venturi situated on the axis of the nozzle between the primary and secondary air swirlers in order to encourage breaking up of the fuel into fine droplets, and ventilation holes opening out close to a tip of the nozzle.

14/ A chamber according to claim 13, wherein the set of pilot head nozzle systems is provided with fairing so as to minimize head losses in the air that flows round the pilot head.