US3019607A

US3019607A - Exhaust shroud cooling

Info

Publication number: US3019607A
Application number: US661383A
Authority: US
Inventors: Dennen J Bunger
Original assignee: Bendix Corp
Current assignee: Bendix Corp
Priority date: 1957-05-24
Filing date: 1957-05-24
Publication date: 1962-02-06
Anticipated expiration: 1979-02-06

Description

6, 1 D. J. BUNGER 3,019,607

EXHAUST SHROUD COOLING Filed May 24, 1957 2 Sheets-Sheet 1 33 NCZZLE +-FUEL 34 IGNITER INVENTOR.

|8 DENNEN J. BUNGER ATToRifiY Feb. 6, 1962 D. J. BUNGER 3,019,607

EXHAUST SHROUD'COOLING Filed May 24, 1957 2 Sheets-Sheet 2 INVENTOR.

DENNEN J. BUNGER m w. ze

ATTORNEY iinitd This invention relates to engine starters and more particularly concerns compact fuel-air combustion starters which have a turbine, driven by combustion gases and geared to drive a turbine aircraft engine.

In certain prior combustion starters, the time of burning was very short, for example, three to four seconds, and it was possible to have a multiplicity of exhaust ports so that heating of the starter 'by exhaust gases was not a major problem. Subsequent developments required that the burning time be longer, and in some cases that one exhaust port in the exhaust shroud be used. With these requirements, it became necessary to prevent the exterior surface of the exhaust shroud of starter from being heated to a temperature above 900 F. Surfaces at this temperature adjacent the aircraft engine are a serious hazard due to the danger of fire in connection with the fuel system. Further, heat conduction to adjacent starter parts from the exhaust shroud was increased.

An object of the present invention is the provision of a starter having means to shield the exhaust shroud and to abstract heat from the turbine exhaust chamber so that the exhaust shroud is maintained at a lower and safe temperature.

An additional object is the provision of a starter having such a shielding and heating extracting means wherein part of the limited supply of air from the starter air bottle is efficiently used for cooling.

A further object is the provision of a starter having an exhaust shroud shield construction including a layer of tubes which is easily constructed and inexpensive.

The disclosed embodiment of the present invention includes means forming an annular exhaust chamber, means in this chamber along the inner surface of the exhaust shroud for shielding the shroud from the exhaust gases and means for cooling this shield with part of the expanded combustion air from the air bottle and for moving the heated air to the combustion chamber.

The accomplishment of the above objects and the advantages of the present invention will be apparent from the following detailed description and the accompanying drawings.

In the drawings:

FIGURE 1 is a schematic, partially-cross-sectioned illustration of a starter embodying the present invention as seen from the top and shows the means for providing the cooling circuit.

FIGURE 2 is a transverse cross-sectional view showing the actual construction and arrangement of the shielding tubes, the exhaust shroud, and the manifolds in side pockets of the shroud. This manifold arrangement is radially displaced in the FIG. 1 schematic for clarity.

FIGURE 3 is a side elevation cross-sectional view of the turbine section of the starter and shows the annular exhaust chamber as actually constructed, rather than schematically as in FIG. 1.

Referring to FIG. 1, it can be seen that conventional reduction gearing 1t and conventional means 11 including a clutch, a jaw advancing mechanism, a starter jaw, and an engine jaw are indicated diagrammatically. Gearing lil and means 11 provide for transmitting torque from the high speed turbine 12 via turbine shaft 13 to an aircraft turbine engine (not shown) connected to the splined shaft 14 which carries the engine jaw. The axial flow impulse type turbine 12 is rotated by combustion gases discharged through a ring of nozzles 15. The exiQQ haust gases from the turbine blades 16 are collected in a generally annularly-shaped exhaust chamber 17 and are discharged through a single oval-shaped exhaust port 18 which is shown as actually constructed and more clearly in FIGURE 2. The inner wall of the annular exhaust chamber 17 is formed by the nozzle ring part of the transverse combustion chamber wall 21 and the asbestos shield or heat barrier 22. It is to be noted that shield 22 extends axially from the transverse support wall 23 to adjacent the turbine 12 and is formed at its outer part to deflect the exhaust gases from the blades 16 radially outward into the annular exhaust chamber and toward the exhaust shroud 24. The short turbine shaft 13 is supported in mounting 25 which includes a seal and bearings. Mounting 25 is carried by support wall 23 which also carries parts of the reduction gearing 10. The actual construction and arrangement of these elements is shown in FIG. 3 and will be further described by use of the same reference numerals.

The combustion gases for driving the turbine are formed in combustion chamber 31 by supplying air to inlets 32 and fuel through nozzle 33 and igniting the mixture by means of igniter 34. The air for the combustion starter is obtained from a small fixed-volume air bottle 35 which is mounted in the aircraft. This bottle has a fixed volume supply of air at very high pressures, for example, about 3000 p.s.i.g. The air from the bottle 35 flows through solenoid shut-off valve 36 and pressure regulator 37 to a small supply chamber 38. The pressure regulator 37 reduces the air pressure from a gradually-diminishing level to about 350 p.s.i.g. The pressureregulated air in chamber 38 is passed through a restricted metering orifice 39 tangentially into an annular chamber 40 which surrounds the fuel nozzle and igniter. Orifice 39 further lowers the pressure of the air. The air from chamber 40 is then passed to combustion chamber through air inlets 32. The swirl chamber 40 provides a collecting chamber for air so that it is uniformly admitted to the air inlets. Air inlets 32 are arranged to introduce air in a predetermined turbulent manner into the combustion chamber 31. Control means (not shown) are provided for automatically opening the air shut-off valve, for simultaneously supplying fuel and for initially energizing the igniter. This control means also includes a turbine speed responsive device which closes the air control valve and stops the supply of fuel after starting an aircraft engine. Such a control system is shown in United States Patent No. 2,742,759, issued April 24, 1956, to Flanigen, et al.

in order to maintain the temperature of the outer wall of the combustion chamber at a lower level during combustion and thereafter, air is passed from the supply chamber 38 which is downstream of the pressure regulator 37 by a conduit 51 to a tubular manifold 52. This air then flows equally through five copper tubes 53 which extend circumferentially along the inner surface of the exhaust wall or shroud 24 to another tubular manifold 5 and then by a conduit 55 having restricted orifice 56 to the annular combustion air chamber 49 which is downstream of the air metering nozzle 39. The showing in FIG. 1 is schematic in respect to the location of the

tubular manifolds

52, 54. FIG. 2 is a showing of the actual location and construction. It is to be noted that the tubular manifolds are adjacent the exhaust port 18 and are in pockets 58 formed in the exhaust shroud 24. The manifolds extend through openings in the transverse wall of the shroud to recesses in the opposite side wall of the exhaust shroud. The arrangement of the tubes 53 and the off-set openings in the

tubular manifolds

52, 54 provide a very satisfactory and inexpensive construction for having the outer part of wall 24 substantially completely shielded from the exhaust gases. It is to be noted that the offset manifolding firmly positions the tubes 53 so that they abut each other and rest against the outer part f the exhaust shroud 24 without being brazed or otherwise attached to the outerwall. Heat conduction is thus minimized. The

manifolds

52, 54 are secured to the vertical side portions of the shroud 24 which forms the pockets 58 (see FIG. 2) by brazing at the ends of the manifolds. The restricted orifice 56 is of such size that the cooling air expands therethrough into the air supply chamber 4% and is of such size that 25-50 percent of the air (a substantial portion) flows through the tubes 53, preferably 30-40 percent is so used.

FIG. 2 shows how the tubes 53 rest against most of the shroud 24 and effectively shield the circumferential part of the shroud. The pockets53 adjacent the exhaust outlot 18 and the off-set manifolding are also shown more clearly.

Referring to FIG. 3, it can be seen that the ring part of the combustion chamber wallZft having the nozzles 15, the outer exhaust Wall or shroud 2d of U-shaped cross section and the outer part of the shield 22 form the annular exhaust chamber 17. Support wall 23 extends beyond the periphery of the turbine 12 and is joined to the exhaust shroud 24 Which in turn, is connected to the outer part of the combustion wall 21. it is apparent that theshield 22 (containing asbestos-type material) deflects the exhaust gases toward the exhaust shroud 24, as above described, and provides a heat barrier for the turbine shaft mounting 25, support wall 23, and the reduction gearing lit). The turbine shaft mounting 25 includes bearing 61 and seal 62. It is to be noted that any heat absorbed by the exhaust shroud 24 will pass by conduction to the housing wall 63 and the transverse support wall 23 and then to the reduction gearing, the turbine shaft mounting and other components of the starter. It is to be noted that the space between combustion chamber wall 21 and the turbine 12 constitutes dead air space. Since-the cold air flowing in the tubes 53 will absorb heat from the swirling exhaust gases, it is apparent that the heat which passes by conduction through the steel parts will be reduced. It is also apparent that tubes 53 interccpt the heat flow to the exhaust shroud, especially the circumferential part, and that cooling air in the tubes will remove the heat from the tubes and also from the part of the shroud which contacts the end tubes.

The operation of the present invention is believed to be readily apparent from the foregoing description. The starter is put into operation by the aforenoted control system so that fuel and air are supplied to the combustionchamber for ignition and burning. The opening of the air control valve 36 also permits cold pressure-reduced air to flow automatically through the tubes 53 at thesame time as air is admitted tothe combustion chamber vvia the metering orifice 39. Thus, the means for maintaining the outer exhaust wall at lower temperatures functions as soon as exhaust gases leave the turbine blades outer wall, this wall is maintained well below 900 F. -A skin temperature of 600700 F. was achieved in tests with the coolant air being heated to 450 F. This heating or" part of the combustion air results in improved cornbustion because the temperature of the total combustion air is increased, after the heated air from tubes 53 is mixed in chamber 40 with the air which is not circulated to the tubes 53. Conduction of heat from the tubes 53 to the wall 24 is minimized because the tubes are not attached to the wall. Since the outer wall 24 is shielded and maintained at lower temperatures, it is apparent that less heat i is transmitted by conduction to the reduction gearing, the housing therefor, and the turbine mounting.

it is to be understood that changes can be made in the disclosed embodiment by persons skilled in the art without departing from the invention as defined by the appended claims.

What is claimed is:

l. A compact combustion starter for aircraft turbine engines comprised of a ring of nozzles; a combustion chamber connected to said nozzles; a turbine arranged to be driven at its periphery by gases from said nozzles; drive means including reduction gearing for engaging an aircraft engine; a short shaft connecting said reduction gearing and said turbine; said shaft being rotatably mounted in a mounting which is positioned between said gearing and said turbine; heat barrier means extending radially outwardly from said mounting between the downstream side of the turbine and a' transverse support with for said reduction gearing; said transverse support wall extending outwardly slightly beyond the periphery of said turbine; an exhaust shroud extending from the outer part of said transverse support wall to a transverse combustion chamber wall; said exhaust shroud having only one exhaust port and a pocket at each side of said exhaust port; said heat barrier means, said exhaustshroud, and said combustion wallbeing arranged so that gases from said turbine are deflected radially outwardly iandso that an annular-like exhaust chamber is formed, wherebylex haust gases are directed radially-outwardly away from the support wall and turbine mounting; cooling-shielding means in said exhaust chamber comprised of a layer of abutting tubesextending circumferentially along and resting on the inner surface of said exhaust shroud from one side to the other side of said exhaust port; said tubes having inlet ends and outlet ends adjacent said single exhaust port; two,longitudinally-extending inlet andoutlet manifolds each having two rows of oifset openings positioned in said pockets in said shroud adjacent said one exhaust port; the inlet and outlet ends of said tub'es' being respectively received in the two rows of said off-set openings of each of said inlet and outlet manifolds;

2. A compact fuel air combustion starter for aircraft turbine engines comprised of a ring of nozzles; a combustion chamber connected to said nozzles; a turbine arranged to be driven at its periphery, by gases-from saidno'zzles;

outwardly slightly beyond the periphery of saidturbine;

an exhaust shroud extending from the outer part of said transverse support wall to a transverse combustion chamber wall; said exhaust shroud having only one exhaust port and a pocket at each side of said exhaust port; said heat barrier means, said exhaust shroud, and said cornbustion wall being arranged so that gases. from, said turbine are deflected radially outwardly and so that anamnular-like exhaust chamber is formed, whereby exhaust gases are directed radially-outwardly away from the support wall and turbine mounting; cooling-shielding means in said exhaust chamber comprised of a layer of abutting tubes extending circumferentially along and resting on the inner surface of said exhaust shroud fromv one side to-the other side of said exhaust port; said tubes having inlet ends and outlet ends adjacent said single exhaust port; two longitudinally-extending inlet and outlet manifolds each having two rows of offset openings positioned in said pockets in said shroud adjacent said one exhaust port;

the inlet and outlet ends of said tubes being respectively received in the two rows of said off-set openings of each of said inlet and outlet manifolds; an air bottle for high 5 pressure air connected by air conduit means to said combustion chamber; said air conduit means including a control valve and a pressure regulator constructed to expand air at very high pressure to air at a decidedly lower pressure; said air conduit means further including a restricted air metering nozzle adjacent said combustion chamber and an air collecting chamber having inlets opening into said combustion chamber; cooling air supply means connected into said air conduit means between said air pressure regulator and said air metering nozzle and extending to said inlet manifold; said cooling air supply means also being connected to said outlet manifold and extending to said air collecting chamber whereby the pressureregulated cold air flows through said tubes when said control valve is opened and then is mixed with air flowing to said combustion chamber; said coolant supply means including a restricted orifice which has a smaller flow path than said air metering nozzle whereby a predetermined substantial portion of air flows through said tubes.

References Cited in the file of this patent UNITED STATES PATENTS